group velocity angle

a b_r (-), mean;

a caret (^), complex

eigen

AbstractThe history of Laminar Flow Control (LFC) from the 1930s through the1990s is reviewed and the current status of the technology is assessed. Early studies related to the natural laminar boundary-layerflow physics, manufacturingtolerances for laminar flow, and insect-contaminationavoidance are discussed.Although most of this publication is about slot-, porous-, and perforated-suctionLFC concept studies in wind tunnel and flight experiments, some mention is madeof thermal LFC. Theoretical and computationalnamics are included for completeness.

the LFC aerody-

(LFC) research that began in the 1930s and flourished

through the early 1960s until it was de-emphasizedbecause of a change in national priorities. During the1970s when the oil embargo by OPEC led to a fuelshortage and high-costfuel, LFC research becameimportant again because of the aerodynamicperformance benefits it could potentially produce for commercial aircraft. The next 20 years of research resultedin numerous significant achievementsin LFC throughwind tunnel and flight experimentsin the United

aircraft are reviewed, including the potential impacts

of insect and ice accumulationon laminar flow extent.From figure 1, two clear eras can be (subjectively)identified over the history of LFC. The first era is theearly wind tunnel and flight experimentsand designtool advancementsin slot-, porous- and perforatedsuction systems through the mid-1960s prior to theOPEC oil embargo, which are covered in section 5.AlthoughmanysuccessfulLFC demonstrationsoccurred in that era, the Vietnam Conflict caused ashift in U.S. national priorities and the demise of the

States and Europe.

major LFC projects.

The balance of this publication presents wind tunnel investigations,

flight research activities, and LFCdesign tool methodologydevelopmentin the UnitedStates and Europe that are devoted to advancing thestate of the art and reducing the risk associated withthe applicationof LFC technologyto subsonic, transonic, and supersonic commercialand military transports.Becausethis publicationis a review,itencompassesmuch of the nearly 60-year history ofLFC research and LFC-relatedresearch to highlightthe many basic flow physics experimentsand theory

Early in the 1970s, the OPEC oil embargo caused

the United States to generate national programs whichfocused on improved aerodynamicefficiencies.Thisfocus reenergizedLFC under the NASA AircraftEnergy Efficiency(ACEE) Program.Many of themajor natural laminar flow (NLF) and LFC projectsunder ACEE demonstratedthe achievement of laminar

developmentwhich have enableddemonstrations.

successful

hardware

Figure 1 and tables 1 through 3 summarize the

LFC projects that are discussed in this overview andhighlight the reference, LFC information, and accomplishment for each project. In section 2, definitionsappropriateto LFC are presentedand the numerousbenefit studies are summarized.In section 3, the manyfundamentalstudies which have led to the currentunderstandingof the flow physics, the manufacturingtolerances necessary for laminar flow, and the designtools

used

to predict

the

extent

of laminar

flow

flow in flight. Sparked by this U.S. success in the NLF

and LFC programs, Bulgubure and Arnal (1992) notedthat laminar flow projects began in France in 1984 togather data that were currently not available in France.Arrospatiale,Dassault Aviation,and a number ofresearchorganizations(includingONERA)wereinvolved in the French program. Then in 1989, theEuropeanLaminarFlowInvestigation(ELFIN)Project was initiated, consisting of four primary elements concentratingon the developmentof laminarflow technology for application to commercialtransport aircraft. These elements were

A transonichybrid laminaron a large-scale

wind

tunnel

flow controlmodel

evaluation(HLFC)

of theconcept

4.

The developmentof a boundary-layersuctiondevice and the development of new wind tunneland flight test techniques for LFCThe developmentof improvedmethods for laminar-to-turbulenttion capability

LFC ar_ discussed

studies

2.1. Definition

computationalflow predic-

A partial-spanflight demonstrationlaminar flow (Birch 1992)

by summarizing

of LFC

of natural

ent flow physics

phenomena;althoughthe samecontrol system may be employed for both problems,the energy requirementsfor relaminarizationcouldtypically be an order of magnitude greater than thatrequired for LFC. Finally, LFC is a capability that isdesigned to benefit an aircraft during cruise by reducing the drag.

An alternate concept

in Harris and Hefner (1987), Wagner et al. (1988),

Wagner et al. (1992), and Hefner (1992). A few bibliographies of LFC are available by Bushnell and Tuttle(1979),Tuttle and Maddalon(1982,1993), andKopkin and Rife (1977). Holmes and Obara (1992)and Holmes, Obara, and Yip (1984) review and focus

Finally, refer to Research in Natural Laminar Flow

and Laminar-FlowControl (NASA CP-2387,1987)and First European Forum on Laminar Flow Technol-

nacelles have overcome earlier design deficiencies. An

active system is usually required to prevent theseboundary-layerinstabilities from causing the laminarflow to become turbulent.

A significant advancementmade in the development ot LFC technologyis the concept of HybridLaminar Flow Control (HLFC). Shown in figure 2,HLFC integrates the concepts of NLF with LFC to

2. Background

LFC, and HLFC are outlined

is referred

boundary-layerinstability termed "crossflowvortexinstability"(discussedin section 3). This instabilitycauses the NLF design to become ineffective and theboundary-layerstate to become turbulent very near thewing le_ding edge. For nacelles, the application of theNLF de dgn has been shown to produce unacceptablelow-spe._d performance;however, some modem NLF

on NLF flight research; Somers (1992) and Pfenninger

to as "natural laminar flow (NLF)." NLF employs a

favorable pressure gradient to delay the transition process. Inherent in practical NLF wings is low sweepand aircraft of small to moderate size. As the wing isswept, _erodynamic performancebenefits are realizedfor hig a-speed aircraft;however,the now threedimensional(3D) flow field becomes vulnerable to a

This overview publication attests to the enormous

amount of research pertaining to NLF and LFC in theliterature. Additional discussions of LFC can be found

sections,

benefit

LFC is an active boundary-layer

flow control(usually suction) technique employed to maintain thelaminar state at chord Reynolds numbers beyond thatwhich is normally characterizedas being transitionalor turbulent in the absence of control. Understandingthis definition is an important first step toward understanding the goals of the technology.Often, a readermistakenly assumes that LFC implies the relaminarization of a turbulent flow state. These are two differ-

According to Mecham (1992), the project team consisted of 24 organizations,

including Deutsche Airbus(project leader), Arrospatiale,Alenia, British Aerospace, CASA, Dassault Aviation,Domier, Fokker,Saab, several smaller companies, six national aeronautical research institutes, and nine universities. Amongthese institutes and universitieswere ONERA, CIRAINTA, DLR, and the Universitiesof Manchester,Bristol, Galway, Lisbon, Lyngby, Darmstadt,Delft,Madrid, and Zaragoza.Section 6 summarizesthemajor U.S. and European LFC programs for the timeframe beginning with the OPEC oil embargo.

In the following

numerous

reduce suction requirements

and reduce system complexity. LFC is complex, involving suction (and ducts,flutes, and pump source) over the whole-wingchord

of NLF,

and the benefits of using

2

(or enginenacelleor tail section).Thekeyfeaturesof

because of the sensitivity of the laminar flow to external and vehicle disturbances(e.g., panel-panel joints,fasteners, access doors). However, drag reduction dueto laminar flow over select portions of a vehicle isachievable. For an aircraft, the wings, engine nacelles,

2. NLF is maintainedover the wing through

propertailoringof thegeometry(pressure)

fuselage nose, and horizontal and vertical tail are candidates for achieving laminar flow. Although the summation of these individualdrag reductionswouldindicate a benefit due to laminar flow (fig. 4), the max-

gross weight(TOGW),operatingemptyweight(OEW), and block fuel (BF) for a given mission, and

dependent;fuel cost,efficiencyamount of(including

the weight of a passenger

for the overall payloadweight). Throughoutthe history of LFC, numerousbenefit studies have been carried out on a host of configurations. The outcome of these studies is describedin this section along with a discussion of the impact offuel cost on LFC benefit.Antonatos (1966) presented a review of the concepts and applicationsof LFC, beginningwith therealizationthat skin friction drag could amount toapproximately75 percent of the total drag for an aircraft. Shown in figure 3, Thibert, Reneaux,andSchmitt (1990) attributedfriction drag to approximately 45 percent of the total drag. Because laminarskin friction can be as much as 90 percent less thanturbulent skin friction at the same Reynolds number,laminar flow would obviously be more desirable thanturbulent flow for reducing the drag of aerodynamicvehicles (except in recovery regions where a severepressure drag penalty can occur because of boundarylayer separation). A vehicle with laminar flow wouldhave much less skin friction drag than a vehicle withturbulent flow. An example of the benefits of laminarflow are shown in figure 4 for a subsonic business jet(Holmes et al. 1985). Unfortunately,achieving laminar flow over the entire configurationis impractical

Lachmann (1961) discussed the design and operational economies of low-drag aircraft, including LFC.This presentationwas one of the few that listed theequations and assumptions of the equations that led toprojected performance.Lachmann noted that the benefits of laminar flow obtained by LFC increased withthe size of the candidate aircraft, with benefits maximized for an all-wing aircraft. Also, if 39 percent ofthe aircraft fuselage could be laminarized for a typicaltrans-Atlanticairline, Lachmann (1961) predicted a10-percent

increase

in L/D.

Chuprun and Cahill (1966) discussed the performance improvements

of aircraft with LFC technologyfrom the systems perspective and noted that the impactof any technology must involve the integrated result ofaerodynamics,structures, propulsion, cost, risk, reliability, schedules (operations),and the sensitivity ofthe proposed concept to the design goals. This integrated result heavily determines the cost-effectivenessof the design concept and whether the technology willbe implemented on the candidate aircraft. When compared with the turbulent baseline aircraft, the important improvementto the aircraft because of LFCwould be an increase in L/D. The amount of improvement would depend on the amount of laminar flowachieved for a given surface geometry and flight condition and the structural weight penalties incurred bythe addition of the pumping system. At a minimum,the benefits of the LFC technology must overcome thepenalties

By notingthat LFC benefitsincreasewithincreased aircraft range, Goethert (1966) demonstratedthe performancebenefit by example. A long-range aircraft designed to carry a payload of 150000 lb some5000 n.mi. could carry the same payload 6250 n.mi.by employing LFC technology,or the LFC aircraftwould be able to carry a reduced payload of 100 000 lbsome 8000 n.mi.

Later, Sturgeon et al. (1976) performed a systems

study to determine the benefits of LFC on long-rangesubsonic transports. Based on a range of 5500 n.mi.and payloads of 200 (52400 lb) and 400 (104 800 lb)passengers,the LFC transport would improve fuelefficiency by 39.4 percent over advanced technologyturbulent aircraft; therefore fuel consumptionwouldbe reduced by 28.2 percent and operating costs by8.4 percent.Pearce (1982) presentedthe benefits of a LFCsubsonic transport compared with an advanced comparable turbulent configuration.The benefits of usingLFC were shown to be consistentwith the resultsalready cited; however, unlike many of the studies,Pearce showed the significanceof both laminar flowextent (i.e., transition location on the wing) and fuelcost. For example, a rise in fuel cost from 45 cents to1 dollar would cause direct operating cost (DOC) to beincreased from 3 to 8 percent with LFC compared withthe turbulent configuration.

tail, and struts. Such an aircraft would

carry 50000kg of payload(or 250 passengers+cargo) and cruise at a Mach number of 0.83. Weaksuction was positioned from 5 to 30 percent chord andit was predicted to achieve laminar flow on about70 percent(HLFC I.

chord

on the upper

surface

of the wing

As illustrated in figure 5, Kirchner (1987) showed

that the benefits of LFC (HLFC and NLF) increasedwith the increased size and range of the candidate airplane. This figure indicates that the benefits of LFC ona long-range subsonic transport could lead to significant fuel savings.Clark, Lange, and Wagner (1990) reported thebenefitsof LFC for advancedmilitary transportaircraft. Based on a 132 500-1b payload transported6500 n.mi. at a Mach number of 0.77, the LFC transport would lead to reductions in TOGW of 4 to 7 percent, fuel weight of 13.4 to 17 percent, and thrust of10.6 to 13 percent and an increase in cruise L/D of18.4 to 19.2 percent compared with the turbulent baseline cor_figuration. The lower and higher values correspondedtolow-wingandhigh-wingHLFCconfigurations,respectively.Arcara, Bartlett, and McCullers (1991) performeda LFC _enefit study for an advanced subsonic, twinengine commercialtransportwith projected1995engine,structure,andaerodynamictechnologyimprovt;mentsinto a HLFC.With laminarflowassumext on 50 percent chord on the upper wing surfaces and horizontal and vertical tails and 40 percenton the engine nacelles, figure 6 shows reductionsinTOGW of 9.9 percent, OEW of 5.7 percent, and BF of18.2 pe.:cent. Additionally,an increase in cruise L/Dof 14.7 percent was achieved compared with that ofthe turbulent baseline. The figure shows the veryimportant location of the suction and resulting laminarflow extent. The analysis included conservativeestimates of the HLFC system weight and engine bleed air(to drive the suction device) requirements.Satisfactionof all operationaland Federal Aviation Regulations(FAR)requirements,such as fuel reservesand

design tool devel-

Supersonic laminar flow control (SLFC) implies

that the test vehicle flies at supersonic Mach numbersand that either LFC or HLFC is employedon thevehicle. Feasibilitystudies by Boeing CommercialAirplaneCompany(Parikhand Nagel 1990) andMcDonnellDouglas Corporation(Poweil, Agrawal,and Lacey 1989) were conductedto determinethebenefits of SLFC applied to the HSCT configuration.The Boeing configurationwas designed to cruise ata Mach number of 2.4 and carry 247 passengers(745 000 lb TOGW) 5000 and 6500 n.mi. The inboardwing was a modifiedairfoilfrom the NACA65A-seriesand had a sweep of 75 (normal Machnumber of 0.62 at cruise), whereas the outboard portion of the wing had a sharp supersonic leading edgewith 47 of sweep (normal Mach number of 1.64 atcruise). The SLFC feasibility study estimated benefitsto be reductions in TOGW of 8.5 percent, in OEW of6.2 percent, and in FB of 12 percent. Thesetook into account the estimated8500-1b

numberssuction-

system weight penalty. The benefits were greater for

an aircraft resized for a range of 6500 n.mi. and areshown in figure 7. With laminar flow covering 40 percent of the wing wetted area, reductions in TOGW,OEW, and FB of 12.6, 9.8, and 16.0 percent, respectively, were projected when compared with the turbulent version of the supersonic aircraft for a range of6500 n.mi. Based on a TOGW of 750 000 lb for theturbulent baseline HSCT aircraft, the projected reduction in TOGW for the laminar aircraft is roughlyequivalent

to the payload fraction

of the aircraft

OEW.

The McDonnellDouglasconfigurationwasdesigned to cruise at a Mach number of 2.2 and carry308 passengers(750000 lb TOGW) 5750 n.mi. Thewing was a cranked arrow wing with most of thesweep at 71 o and the outboard 30 percent span of thewing swept 61.5 . The SLFC feasibilitystudy forapplication to the HSCT found reductions in TOGWof 8 percent and FB of 15 percent and an increase incruise L/D of 15 percent. Whereas the Boeing conceptemployed a leading-edgesuctionspanwise suction strip at about 40McDonnellDouglas concept hadsuction and a continuous low levelthe control surfaces.

strip and a second

percent chord, thelarge leading-edgeof suction back to

Based on limited supersonic data, Kirchner (1987)

showed in our figure 8 that an increase of 10 to

30percentin L/Dsupersonic

is expected by using SLFC on the

high-speed civil transport. Pfenninger and

Vemuru (1988) presented a strut-braced, highly swept

wing SLFC long-rangetransport design which wascapable of acquiring values of L/D of 19 to 27 at aMach number of 2 and 16 to 22 at a Mach numberof 2.5.Aerodynamicperformancebenefitsbought byskin friction drag reduction can translate into reducedoperating costs of an aircraft. Figure 9 shows the jetfuel cost per gallon and jet fuel as a percentage of thecash operating cost for the industry over some 20 yr.From these data (Anon. 1985, 1995a), the criticaltimes in the industry are evident when fuel costs grewin the late 1970s and early 1980s and briefly in the1990s. The rapid increase in fuel cost in the 1970sinspired the drag reduction program in the UnitedStates, including NLF and LFC flight test programs. Inthe 1990s the cost of fuel has become a small fractionof the operating cost for the industry and, therefore,the demand for technologiessuch as LFC have diminished. However, similar to the OPEC oil embargo inthe early 1970s that led to a diminished supply of fueland subsequent rise in prices (large demand and lowsupply), technologistsin the government laboratoriesand in industry must be poised to cope with futureuncertaintyin fuel cost (one of many external influences on the demand for innovation). Note, that the

3. LaminarMethodology

Flow

Control

Design

For a LFC design (a wing, for example), the analysis begins by defining an initial wing geometry. Withwing geometry defined, the wing pressures and velocities can be obtained by using transonic wing theoryand/or ComputationalFluid Dynamics(CFD). Theinverse approach of prescribing a target pressure distribution and solving for the wing geometry is thenused. After obtaining the external flow field for thefinal geometry,boundary-layerand stability theorycalculations are used for determining the suction flowrates and distributionfor the desired transition locations. With the suction flow rate determinedfromboundary-layerstability considerations,the pressuredrop through the skin must be set to obtain a reasonable subsurface compartmentationscheme and perforation spacing distributionfor the desired suctiondistribution.The process is iterative until an acceptable design is obtained. Finally, the suction systemducting and compressor specificationsare prescribed.Other key issues, covered in this section, that mustbe understood

for LFC design are

1. The physics associated with the laminar to turbulent boundary-layer

transition process

rise in fuel price in the early 1990s was spawned by

the Iraq invasion of Kuwait. The yearly consumptionof $10.5 billion in 1981 has only dropped to $7.7 billion in 1994, which reflects a reduction in fuel cost

Impact of surface tolerances--roughness,

and an increase in fuel consumption.

In summary, LFC can lead to reduced skin frictiondrag and thereby reduced fuel consumption. This benefit can lead to either an extension in range for thesame aircraft or to reduced aircraft weight for a fixedrange. For the latter case, less engine power is requiredand reduced emissions, noise, and operating costs canbe expected from the LFC aircraft. Noise and emissionreductions have become ever more important and global pollution becomes an important variable in thedesign concepts of the future. Although fuel cost hasdecreased in recent years, the total volume of fuel consumption has increased and the potential fuel savingsdue to LFC remain a significant cost savings to theindustry.

bance, a dynamic instability,

termed the crossflow(CF) disturbance, is an important factor in the extentof laminar flow realized. The presence of TS and CFdisturbances in the boundary-layerflow is dependent

The first major theoreticalcontributionsto the

study of boundary-layertransitionwere made byand Rayleigh(1879, 1880, 1887). Althoughtheseearly investigationsneglected the effects of viscosity,the second derivative of the mean velocity proved to

on the pressure gradient and on the wing sweep angle.

As shown by Gray (1952), Anscombe and Illingworth(1956), and Boltz, Kenyon, and Allen (1960) forswept wings and by Gregory,Stuart, and Walker(1955) and Reilly and Pfenninger (1955) for rotatingdisk flow, CF disturbancesare characterizedby coro-

be of key importancein explainingboundary-layerinstabilities. These fundamentalstudies proved to bethe basis for future breakthroughsin theoretical devel-

tating vortices (sketched

in fig. 12). For example,Anscombe and Illingworth (1956) used a symmetricairfoil with a 4-ft chord in a wind tunnel experiment to

opment, including inviscid jet-flow

instabilitiesandshear-layerinstabilities.Adding viscous effects, Orr(1907) and Sommerfeld(1908) developed an ordinarydifferential equation (Orr-Sommerfeldequation) thatgoverns the linear instability of two-dimensionaldisturbances in incompressibleboundary-layerflow on

study the flow on the wing swept from 0 to 50 . The

results showed that at angles above 25 to 30 , a critical speed could be found which led to "striations"inthe surface flow visualizationwith transition between

flat plates. Later, Squire (1933) accounted for threedimensional waves by introducinga transformationfrom three to two dimensions. This analysis showedthat two-dimensionalwaves were dominant in flat-

effect of sweep and Reynolds number on transition is

shown in figure 13 (Anscombe and Illingworth 1956).The figure serves to provide a visual qualitative influence of wing sweep. They further noted that as thetransition front moved forward, the laminar boundary

50 and 60 percent chord. As the speed of the free

layer became more sensitive to surface conditions and

the number of turbulent wedges increased. This sensitivity was a unit Reynolds number influence; wherebythe critical height of a roughnesselement affectingtransition decreasedwith increase in unit Reynolds

Schlichting(TS) instabilities,and Liepmann(1943)and Schubauer and Skramstad (1947) experimentallyconfirmed the existence and amplificationof these TSinstabilities in the boundary layer. One can visualizethis disturbanceby rememberingthe image of waterwaves created by dropping a pebble into a still lake or

number (discussed

in section 3.2).

At the same time, Gray (1952) investigated

theeffect of wing sweep in flight using the ArmstrongWhitworthAW.52aircraft.Visualizationwas

puddle. In this image, the waves which are generated

decay as they travel from the source. Such is the casein boundary-layerflow, except that the waves willgrow in strength when certain critical flow parameters(say Reynolds number) are reached and lead to turbulent flow.

achieved through sublimation,

or liquid evaporationfrom china clay techniques. Most of the results are forsweep angles of 25 to 50 , chord locations from 3 to17 ft, and speeds from 50 to 500 knots at an altitude of40 000 ft. Additionally,a Meteor Fin with 25 sweep,a Sabre F.86 with 39 wing sweep, an Avro 707ADelta, and a Hawker P1052 were also tested. Gray

Taylor-Grrtlervortex disturbancesarise when thesurface geometry becomes concave and are reminiscent of counterrotatingvortices. A sketch of thisvortex-disturbancestructure is shown in figure 11.

(1952) concluded that the leading-edge

radius was adirect measure of the limit of laminar flow for all modem flight speeds for sweep angles more than 20 or25 . The amount of laminar flow decreaseswith

The design engineer would have to be sensitive to this

disturbance only if there is concave curvature such ason the lower surface of some wings; otherwise, thisdisturbance is not too significant for LFC applications.See Smith (1955), Wortmann (1969), and Hall (1983)for moredetaileddiscussionsof Taylor-Grrtlerdisturbances.

increased

leading-edge

radius.

Similar

to the results

presented by Anscombeand Illingworth(1956), theresults of Gray (1952) showed that for a given sweepangle, laminar flow was lost as the speed is increasedto a critical speed. Since those early experiments,7

ble modes were observed

(1967),

modes

caused

below

Re 0 = 230.

This tbaee-dimensionalbase flow was a similaritysolution of the Navier-Stokesequations; hence, its useis advantageousin stability analyses. With a nonparallel theory, Hall, Malik, and Poll (1984) determinedneutral :urves with and without steady suction andblowing and demonstratedthat the attachment-line

laminar flow.

by Gaster

these

Accountingfor all linear terms and using aneigenva_ue-problemapproach,Hall, Malik, and Poll(1984) .'tudied the linear stability of the attachmentline boundary-layerflow called swept Hiemenz flow.

boundarlayer can theoreticallysmall amounts of suction. The

Transition along the attachment line can be prevented by designing

the attachment-lineReynoldsnumber not to exceed some critical value. This wasout in experiments

of theory;

line of swept wings led Poll (1979, 1980) to perform

additional experimentswith the swept circular modelof Cumpsty and Head (1969). Like PfenningerandBacon (1969), Poll observed disturbances that amplified along the attachment line. He noted that no unsta-

to prevent turbulent attachment-line

contaminationareillustrated in figure 16 (from Maddalon and Braslow1990). For LFC or HLFC, strong suction can also beused at the fuselage-wingjuncture to relaminarize theflow, and mild suction can be used thereafter on the

drawn

modes

transition to occur at about Re 0 = 240. A continued

interest in the transition initiated near the attachment

portion

of the whole wing would have turbulent flow. Clearly,

this can be understood by viewing the illustration infigure 15 (from Wentz, Ahmed, and Nyenhuis 1985)for the attachment-lineregion of a swept wing. Turbulence (or attachment-linecontamination)from thefuselage boundary-layerflow can sweep out onto theattachmentline and cause the entire wing to beengulfedin turbulent flow. However,a turbulencediverter such as Gaster's bump (Gaster 1965) can beeffectively used to establish a laminar attachment line;this allows the potential for continued laminar flow onthe attachment line. Some methods which can be used

leading edge to maintain

the boundary-layerinstabilities,theylaminarflow was stable to small-

amplitude disturbances up to Re 0 = 245 (which corresponds to the top speed of the tunnel). At the same

which lead

line, the outboard

recorded oscillationshad preferred frequency bandsthat changed with tunnel speed and that this behaviorwas reminiscent of traveling-waveinstabilities. Fromhis measurements,he concludedthat the small-

cially trippingobservedthat

to transitionover the wing chord, attachment-lineinstabilities are possible and can be correlated for natural transition in the linear limit with the Reynoldsnumber of the flow. If transition were to occur at somelocation

in the flow by a hot-film

line. He noted that the

Reynolds numbers Re 0 below 170. Later, Cumpsty

and Head (1969)experimentallystudiedlargeamplitude disturbancesand turbulent flow along theattachment line of a swept-wing model. Without artifi-

wing (or nacelle, etc.). For large sweep angles, LFC or

HLFC suction is used in the leading-edgeregion tosuppressthe normallyrapid growthof the CFdisturbances,and then the pressureon the wingsurface is tailoredto minimize the growth of alldisturbances.to TS and CF disturbances

to cause this wedge to cling to the fuselage

possible;thereby,laminar flow would occur in aregion close to the fuselage. The author knows of nostudy which has investigatedthe potential instabilityof the interface between a turbulent wedge and laminar flow over a wing; however, Hilton (1955) has usedthe concept of tailoring the streamlines to the fuselageto obtain a drag reduction.

studying leading-edge contamination,

Pfenninger (1965) discovered

throughments that laminarflow could be

as much as

flight experiobtainedfor

Re 0 < 100 but leading-edge

contaminationoccurredfor Re 0 > 100. Gregory and Love (1965) found thatcomplete turbulence occurred for Re 0 > 95 in theirwind tunnel experimentson a swept airfoil. Flightexperiments by Gaster (1967) showed that turbulentspots were first observed for Re 0 > 88. Cumpsty andHead (1969) and later Poll (1985) used a swept modelin a wind tunnel to show that turbulence was dampedfor Re 0 < 99 and the leading edge was fully turbulentfor Re 0 > 114. Arnal, Juillen, and Casalis (1992) useda swept-wing model in a wind tunnel to show thatleading-edgecontaminationwasobservedat

In summary, for wing sweeps from 0 to 10 , TS

disturbancesamplify and cause natural transition. Ifthe design pressure gradient is favorable (acceleratingflow), longer runs of laminar flow can be realizedbecause the TS-disturbancegrowth rate is suppressed,whereas the opposite is true with an adverse pressuregradient. Wing design should minimize the growth ofthese disturbances to enable long runs of laminar flow.Between wing sweep angles of 10 and 30 , both TSand CF disturbancesare present, amplify, and causetransition; much of the flow physics associated withthe nonlinear interaction of these modes is unknown.For wings swept greater than 30 , CF disturbancesdominate, amplify, and cause transition--oftenverynear the leading edge of the wing. Hence, LFC isrequired to achieve laminar flow on highly sweptwings. Also, the leading-edge radius affects the stability limits of flow along the attachmentline, withincreased leading-edgeradius being destabilizingtothe flow.

critical Re 0 care must be taken so that the flow is not

tripped. Wind tunnel experiments by Carlson (1964)indicatedthat the Reynoldsnumberbasedonboundary-layermomentumthickness at the front of

Tolerances

for Laminar

Flow

Roughness,waviness, steps, and gaps are issuesrelated to manufacturingtolerances.Joints, rivets,screw heads, and panel jointscontributeto theroughness-steps-gapsissue, and stiffness of the skinwith imposed loads and overall manufacturedskinsmoothnessare ingredientsin the wavinessissue.

the attachmentline should be Re 0 < 150 for verysmall disturbancesand Re 0 < 100 for large disturbances. As many flight experimentshave shown,maintaining NLF on the attachment line is possible,and the momentum-thicknessReynolds number can belowered by reducing the leading-edgeradius or unitReynolds number. Decreasing the leading-edge radiushas the compoundedbenefit of decreasing the chordwise extent of the crossflow region and providing amore rapid acceleration of the flow over the wing.

Additionally, a turbulent wedge, originating at the

fuselage-wing-leading-edgejuncture, can sweep outover a portion of the wing root region and is a concernfor NLF and LFC wing design. Clearly, one wouldattempt to optimize the fuselage-wingjuncture point

an overview9

of flight

test

experiments

conducted

(mainly)at CambridgeUniversityin(1938)

stated

England. Jonesthat the main conclusionsfrom those

flight experiments.

3.

In experimentsto examine transition in flight,Stephens and Haslam (1938) used a Hart K1442 aircraft which had a 2D wing test section and a SnarkL6103 aircraft which had a mildly swept-wingtestsection. Among the reported results, spanwise ridgesof height 0.002 in. caused transition to move forwardat chord Reynolds numbers of 5 106 and more; the

were

Drag predictionsfor moderatelythick wingshapes can be made based on smooth flat-plateskin friction data if the transition points wereknown for the wing

Surface roughness flight experiments described by

Bicknell (1939) were conducted on a Northrop A- 17Asingle-engineattack airplane. The focus of the studywas to characterizethe impactof conventionalmanufacturer-inducedroughness and gaps (rivets, lapjoints, access panels, and hinges) on drag. The resultsfor a standard wing were comparedwith a smoothwing at a chord Reynolds number up to 15 x 106. The

transi-

The flight and wind tunnel tests have provided our

current understandingof the mechanismswhich causetransition to move forward because of surface imperfections. The impact of a surface imperfection (such asa rivet head) on the transition location can be viewed

wing was made smooth by filling lap joints and

cementingpieces of rubber sheeting to build up theareas of rivet protuberance.The results show that a50-percent increase in the profile-drag coefficient wasobtainedsmoother

either by looking at the transition location as a function of imperfection

size for a fixed unit Reynoldsnumber or by keeping the size of the imperfectionfixed and looking at transition location as a function ofunit Reynolds number. The illustration in figure 17(Holmes et al. 1985) depicts the latter case, where theamount of laminar flow is decreasedas Reynoldsnumber is increased. The problem is then to determinewhat roughness height and shape for a given Reynoldsnumber will cause a reduction in the amount of lami-

with thewing.

rough

wing

compared

with

the

At the Royal Aircraft Establishment

(RAE) inEngland, Young, Serby, and Morris (1939) reportedon the impact of camouflage paint, snap rivets, flushrivets, lap joints, and leading-edgeslats on wing dragof the p,ototype Battle. The Battle had wings with lowsweep, with each wing containing three bomb doorson the t nderside of the wings (reason for joint study).The tesls were conducted by fitting specially preparedskins over portions of the wings (approximately,NACA 2417 airfoils). The range of chord Reynoldsnumber was 12 106 to 18 106 with approximateunit Reynoldsnumbers per foot of 1.2 x 106 to1.8 If -6. Both the drag due to the variation of transi-

nar flow obtainable. In either case, the imperfection

stimulates eigenmodes in the boundary layer; the linear stability of the flow dictates whether these modeswill grow or decay as they evolve in the flow. However, as the height of the imperfectionor unitReynolds number increases, a point is reached whenflow separation occurs because of the surface imperfection. At this point, inviscid instability arising fromthe inflectional velocity profile can grow and inducetransition. Or if the imperfectionis sufficiently large,linear instability amplificationis "bypassed"and transition follows by way of a nonlinear process. Our current understandingof imperfectionssuggeststhatlarger criticalstep heights can be realizedwithrounded steps because a reduced region of separationand reduced inflectional instability growth are encountered in the experiments.

tion location (due to protuberance) and drag due to the

protube'anceitself were measured in the course of theflight test. For the Reynolds number per foot of1.8 1C_6,transition was forced upstream of the protuberance of interest. In brief, the conclusionsof thisflight te_t were.

Camouflage

paint did not infuence

the transi-

tion points; however, painting the wings of the

Battle-typeaircraft reduced its top speed byabout 3 to 4 percent

10

lift coefficient

Span rivets both increased drag and affected

the transition point; for example, completelyfastening the wings of the Battle-type aircraftwith rivets 0.04 in. high and 0.25 in. widecaused a decrease in the top speed of the aircraft by about 2.5 percent

and Reynolds

number

(i.e., quantitative

evaluation was not possible), some qualitative comparisons can be made with reference to surface andengine conditions. A two-tube rack was used to measure the transition location. For the design lift coeffi-

cient (C L = 0.2) and Reynolds number of 26.7 x 106,

tion point was affected with this type of rivet;

the implementationof flush rivets should be asfar back from the leading edge of the wing as

is at approximately45 percent chord. For this bestlaminar flow case, the surface had a waviness amplitude of 0.001 in., which was obtained through polish-

possible

ing the surface. For the same flight conditions and a

surface waviness amplitudeof 0.005 in., transitionoccurred at 32.5 percent chord. This early work gavean indication of the influence of waviness on laminar

Flush rivet drag was negligible

Ten unchamferedrearward facing lap joints(1/16 in. high) decreased the top speed of theBattle-typeaircraft by 2.5 percent; however,chamfered to a gradient of 1:5 led to only a1.5-percent speed reduction

flow extent; however,

ferences in flight test results and wind tunnel results

were directly impacted by residual turbulence, even inthe "quiet tunnels" of that time.

The addition of a leading-edge

slat to half thewing of an aircraft with transition occurringnear the leading edge led to a top speed reduction of 1 to 2 percent for a very well-fit slat andof about 2.5 percent for an average-fit slat; iftransition was not at the leading-edgeregion,then the slat-incurreddrag would be greaterthan if it were

Fage (1943) performed the first systematic wind

tunnel experiment to characterize the surface wavinessimpact on laminar flow (point of transition) for a flatplate boundary-layerflow. The experimentswere carried out using "corrugations"--smoothbulges andhollows and flat ridges---on one side of a smooth flataluminum plate which had an elliptical leading edge.Although the tunnel could produce sufficiently cleanflows up to a tunnel speed of 140 fps, the experimentswere carried out so that the corrugations impact transition well below 140 fps and are not affected by freestream turbulence in the wind tunnel. Positioned 20 in.

Formulas for estimating the drag effects due to

rivets and lap joints were shown to be in goodagreement with experimentalresults; althoughthe formulas for describing the drag due to rivets and lap joints are very important for turbulent configurations,the capability to predict theimpact of the protuberanceon the transitionlocation is more significant for NLF and LFCapplications

Wetmore,

Zalovcik,

and Platt (1941)

performed

because no surface wavelengths

were presented, the flight data cannot be used for waviness correlations. Finally, it was recognized that dif-

downstream of the leading edge, a strip of spring steel

was used to form bulges and hollows and a piano wirewas used for ridges. Small surface tubes (mounted onthe plate) were used to indicate when a corrugationcaused transition to move forward as the tunnel speedwas varied.

flight investigationto study the boundary-layercharacteristics and profile drag of a 2D laminar flow airfoilat high Reynolds numbers. They used a Douglas B-18aircraft modified with an NACA 35-215, 17-ft chord

For this zero pressure gradient case, Fage (1943)

found empirical expressionswhich gave an estimatefor the minimum height of spanwise bulges, hollows,and ridges that affects the position of transition in theexperiments.The experimentsshowed that the minimum height is not especially dependent on the form ofthe corrugation, and it appeared that the flow conditions that impact the transition location were related tothe local separation of the laminar boundary layer.

by 10-ft span test panel positioned on the wing 13 in.

outboard of the propeller-pulledengine of the aircraft.The test covered Reynolds numbers from 20 x 106 to30 x 106 and included variations in power and surfaceconditions.Engine power variationswere made todetermine the impact of the engines on profile drag.Although there was no fixed relationship between the11

However,asFagenoted,it wasnotsimple relations

expected that these

take into considerationall flow condi-

ness b) the use of appropriate filler and careful rubbing down

the surface.Surfacewavinesswasmeasured to be less than 0.001 in. The results showed

tions. In particular, only flow separation was considered and the stability of the flow downstream of the

a 26-percent decrease in the drag coefficient compared

with previous flight test results. Laminar flow wasrealized to between 50 and 60 percent chord of the testsection (the pressure minimum was designed for about50 percent chord). The conclusionsfrom this flighttest were in agreementwith the previous King Cobratest; namely,reducingthe surfacewavinessto0.001 in. led to significant runs of laminar flow forflight Reynolds numbers in the range of 20 x l06.

corrugationshould be accounted for as well. Fage'swork did not include the effects of compressibilityorsweep.At the same time, Braslow (1944) was studyingthe impact of roughness on transition in a less systematic manner than Fage (1943). The effect of variouscamouflage paints and the painting procedures on thedrag characteristicson an NACA65-420 airfoilsection were examined.Using the LangleyLowTurbulencePressure Tunnel (LTPT), Braslow (1944)showed that a carefully applied camouflagepaintedsurface could retain the low-drag characteristicsof theairfoil up to chord Reynolds numbers of 22 x 106. Thismaximum Reynolds number could not be overcome

The earlier wind tunnel and flight experiments

served to illustrate the impact of surface smoothness(roughness and waviness) by demonstration.The following subsections present the current understandingof surface smoothness,building upon these earliertests.

unless some light sanding was applied to the painted

finish.This experimentdemonstratedthe impactroughness could have on drag (or transition) with unitReynolds number variation.

3.2.1. Waviness

Carmichael,Whites,and Pfenninger(1957),Carmichael(1959), and Carmichaeland Pfenninger(1959) developedthe basis for "allowablewavinesscriteria' for swept and unswept wing surfaces, influenced by compressibility,suction, single bulges, multiple waves, and wing sweep. The criteria are stillvalid today and were based on the available flight testobservations.Flight test experiments were carried outby using the F-94A airplane with 69 suction slots asdescribed by Groth et al. (1957). Sinusoidalwaveswere obtained over the width of the test section byapplying paint with the wavelength specified by masking tapt.. Wave height and length were varied in aregion of growth(28 percent chord) prior to thesuction influencingthe disturbanceevolution.Theresults :;howed that the extent of laminar flow was

Smith and Higton (1945) reported the results of

King Cobra flight tests to determine the (surface) criteria for laminar flow and the practicality of meetingthe necessary requirements.The impact of rain, dust,insects, and surface-finishpolish on the flow wasassessed.Dust and water accumulationdid notincrease the measured drag, whereas as the temperatures increased in April 1945, it became impossible tofly without insect contaminationaffecting drag measured in flight. Also, the results showed that reducingthe waviness to _+0.001 in. led to runs of laminar flowto 60 to 65 percent chord. Gray and Fullam (1950)reported wind tunnel tests for the King Cobra wingmodel in the RAE No. 2 11.5- by 8-foot tunnel. Consistent with the flight experiments, low drag was realized for Reynolds numbers of 15 x 106; however, theexistence of turbulence in the wind tunnel, which isnot present in free flight, caused some degradation of

The allowable height for aft-facing steps is one half

the allowable for forward-facingsteps. The allowablegaps for flow over the gap is(Re/ft)g

= 15 000

(3)

and the allowable gap width for flow along the gap isone seventh the gap width for flow across the gap.3.2.3. Three-Dimensional

Surface

Discontinuities

The flow tolerance to roughness was also investigated in the flight test. Single and multiple sphericalshaped glass beads and steel disks were used as roughness on the test section. At 22 percent chord, criticalroughnessheights of 0.0105, 0.007, and 0.0055 in.were obtainedfor a single sphere, a single disk(Height/Diameter= 0.167), and a multibead band ofdistributed roughness, respectively,for a Mach number of 0.68 and altitude of 26 000 ft. At 2.5 percentchord, where the boundary layer was much thinner,the critical heights decreased to 0.007 in. for a singlesphere and to 0.004 in. for the single disk for the sameMach number and altitude. Carmichael,Whites, andPfenninger (1957) exploredcal roughness condition

the definition

of the criti-

U/ka = (59)_

O00_____c_ccos2A11/2

_,Re_'2

Criterion

for two-dimensional

ties can be found

in Braslow

surface discontinui-

Braslow et al. (1990). The allowable

forward-facing steps is(Re/ft)h

= 1800

(4b)

local kinematic viscosity, and U/ is the local potential

velocity. Equation (4a) should be used to determinecritical roughness heights near the leading edge of thewing, and equation (4b) should be used in other thanthe leading-edgeregion. Essentially,the flight testresults showed that these parameterswere a linearfunction of the roughness height.

Discontinuities

and Fischer

UkkRe k = -vk

where k is the height of the roughness, Uk is the local

velocity at the top of the roughness particle, v k is the

doubling the spanwise criteria (Braslow et al. 1990).

An example of the waviness criteria for a LFC airplane with a given wing sweep, Mach number, andaltitude is shown in figure 18 (from Braslow andFischer 1985).Surface

(4a)

(1)

For multiple waves parallel to the wing span, the critical waviness becomes one third of the single-wave criteria, and chordwisewave criteriaare found by

3.2.2. Two-Dimensional

Re K -

(1985)

Braslow and Knox (1958) proposed a method for

determiningthe critical height of three-dimensionalroughnessparticleswhich would cause prematurelaminar-to-turbulentboundary-layertransition.Anequation was derived which related the critical roughness height to local flow conditions (i.e., the local

and

step height h for

(2)13

temperatureandvelocityconditionsin the boundarylayer).The resultswerepresentedfor zero-pressuregradientflow for Machnumbersfrom0 to 5.A roughnessReynoldsnumberRek of between250and600for Machnumbersup to 2 apparentlycausedpremature transition.Then, basedon the assumedRekwhich

Braslow and Maddalon

(1993, 1994) discussedroughness-relatedresults of the Jetstar LFC flight test.A ratio of roughness diameter to height between 0.5and 5.0 is permissiblein the high crossflow region ofswept-wing flow.

3.3. Critical

An example of the critical roughness height with

altitude for a fixed Mach number is shown in figure 19(Braslow and Fischer 1985). As the altitude increases,the unit Reynolds number decreases and the allowablecritical roughness heights can therefore increase.

with

includes

gradient

Depending

on the relationship

between

The criticalwave heightincreased number of waves

decreases

As part of the Saric (1985) review of LFC control

stiction for AGARD,the issue of transition

pressure-taprecords, and the state of the boundarylayer was determined to be either laminar or turbulentwith a :;tethoscope. The critical suction was affected

surface

wavelengthand the disturbance(TS or CF),transition can move forward or be postponed inthe CF-disturbanceregions (due to wave superposition and relative wave phase)

for LFC

used to design suction through isolated holes or a row

of holes. Nondimensionalparameters were determinedfrom rc suits over a large range of boundary-layerReynolds numbers, tube velocities, and hole configurations.Tubevelocitiesweredeterminedfrom

induced by

the surface imperfection could cause the amplification of TS disturbances, which would causepremature transition.

Parameters

The earliest fundamental understanding

of the critical suction issue was reported by Goldsmith (1955,1957), where experimentswere conductedin theNorthrop 2-in-diameterlaminar flow tube to determine universal critical suction curves that would be

A local separationregion due to the surfaceimperfectioncould cause Rayleigh'sinflectional instability, which could cause transitionto move forwardThe local adverse pressure

Suction

caused by local streamwise vorticity generated in the

boundary-layerflow over a suction hole was brieflycovered. Essentially, the threshold parameters are notknown when these vortices appear nor what strengthand impact they have on the flow instabilities. Theseparametersinvolve hole size, suction flow rate, holespacing and geometry, and hole inclination.

The currentunderstandingof the mechanismswhich cause transition to move forward due to surfaceimperfections

Th,'se wind tunnel and flight experiments demonstrated the sensitivity of the flow to the surface definition. They also showed that with some careful surfacepreparation,laminar flow could be obtainable.Thestringentsurface smoothnessand waviness criteria(tolerances) for laminar flow posed a major challengefor research in the 1950s and 1960s. A partial explanation for the descope of subsonic LFC in the 1950s wasattributable to the severe surface manufacturingtolerances required to achieve laminar flow. Howeverthe manufacturingtechnologiesof the 1990s havematured to the point that surface definition tolerancesare more readily achievable.

by the hole diameter, the hole spacing, and the

boundaJy-layerthickness. The suction was adjustedfrom a flow condition which was turbulent until theflow be_:ame laminar. This suction level which led tolaminar flow was called the critical maximum value.

with

When a sufficiently low suction level is reached that

any further decrease in the suction would lead to aturbulentflow, the minimumsuction values wereobtained. The critical maximum suction arose because

multiple spanwise stringers, and countless fasteners

(e.g., rivets) on the surface prevented achieving laminar flow. On research aircraft, fillers were used eveninto the 1980s to smooth problem areas of the surface.With the advent of bonded sandwichconstructionmethods, the productionsurface became as good asthe production mold definition. The surface structurebecame sufficiently stiff so that adequate waviness criteria could be maintained under loads (in subsonic andtransonic aircraft) and the new production capabilityin the 1990s has solved (in principle) the task of manufacturing laminar flow quality surfaces.3.4.1. JointsPotential

issues still remain

associated

with struc-

tural joints. The issue of critical waviness caused by

these intersectionsmust be part of the design process.The intersectionof these major structures may havefasteners which protrude above the surface and causeflow interruptionby way of steps and gaps. To avoidthis problem, a recessed intersection region could beemployed, which would remove the fastener issue andcould require a flush-fillrecessed connectionarea.

techniqueto cover theSimilar to the structural

joints issue, access doors are a normal feature on aircraft and require special attention for laminar flow tobe achit;vable. Flush mounting to within a few thousands of an inch is required; sealing the access panel isalso reqaired to prevent air bleed from the panel.3.4.2. HolesCorlparablewith early analysis on the Jetstar(Powell 1987), Boeing 757, and F-16XL (Norris 1994)LFC flight test articles, Parikh et al. (1990) studied thesuction system requirements(based on computationalanalysis) for SLFC on a High-Speed Civil Transport(HSCT). In the analysis, the perforated skin had holediameters on the order of 0.002 in. (0.05 mm), holespacing of 0.01 in. (0.025 mm), and a skin thickness of0.04 in. (0.1 mm). With this information,it is clearthat millions to billions of holes are required for alarge-scale wing. For example, the hole spacing suggests that 10 000 holes are contained in a square inchor 1.4 million holes are found in a square foot. Fora large applicationsuch as the proposedHSCT,

The improvementsin aerodynamicefficiencydirectlyscale with the amountof laminarflowachieved. Hence, the designer must be able to accurately predict the location of boundary-layertransitionon complex, three-dimensionalgeometries as a function of suction distribution and suction level (or theaccurate predictionof the suction distributionfor agiven target transitionlocation).Pressure gradient,surface curvature and deformation,wall temperature,wall mass transfer, and unit Reynolds number areknown to influence the stability of the boundary layerand transition location. For practical HLFC designs, itis imperativeto be able to accuratelypredict therequired amount, location, and distributionof wallsuction (or thermal control or any other control technique) to attainlocation.

3.5.2. C1 and C2 Criteria

At ONERA, Arnal, Juillen,performedN-factor correlations

and Casaliswith wind

(1991)tunnel

experimentalresults of a LFC suction infinite sweptwing. The motivation for the study was to gain fundamental understandingof the transition process withsuction and to test the methodologiesdevelopedatONERA-CERTfor three-dimensionalflows.Thestreamwise instability criteria were based on an extension of Granville (1953). Two crossflow transition cri-

3.5.1. Granville

Criterion

teria have been developed by Areal, Habiballah,

andCoustois (1984) at ONERA and are referred to as C1and C2. The C1 criterion involves a correlationof

Granville (1953) reported on a procedure for calculating viscous drag on bodies of revolution anddeveloped an empirical criterion for locating the transition location associated with low-turbulenceflows.

transitiononset integralvalues of the crossflowReynolds number and the streamwiseshape factor.The C2 criterion is a correlationof transition onsetwith a Reynolds number computed in the direction ofthe most unstable wave, the streamwise shape factor,and the free-stream turbulence level. The results dem-

Low (or zero) turbulence characteristic

of flight orlow-turbulencewind tunnels and high turbulence characteristic of most wind tunnels are the two problemsconsidered relative to a transition criterion. The lowturbulence case assumed that transition was TS distur-

onstrate that the transition criteria cannot be applied in

regions where the pressure gradient is mild becausethere is a large range of unstable directions.In thatregion, one cannot look only at pure streamwiseorcrossflowinstabilities.The C1 criterion gives badresults with wall suction present; however, the C2 criteflon correctly accounts for wall suction.

(1936), Hall and Hislop (1938), Schubauer

and Schubauerand Skramstad(1948),

3.5.3. inear Stability

Granville (1953) showed that a variety of flight and

low-turbulencewind tunnel data collapsed into a crite-

The equations governing the linear stability of disturbances in boundary layers were first described byOrr (1907), Sommerfeld(1908), and Squire (1933).These equations are ordinary differentialequationsand are referred to as the "Orr-Sommerfeldand Squireequations." Although the growth or decay of smallamplitu:le disturbancesin a viscous boundary layercould b,_ predicted by the Orr-Sommerfeldand Squireequations(within the quasi-parallelapproximation),the ability to predict transition came in the 1950s withthe semi-empiricalmethod by Smith (1953). Thistransition-predictionmethod---callede N or N-factor

rion (curve) based on Re0, T - Re0,N, which is the difference between the momentumthickness Reynoldsnumber

at transition

and at the neutral

point, versus

02 d_v d._'eter.

which is the average pressure

This

correlation

was

gradient

demonstrated

paramfor two-

dimensional flows and is shown in figure 21 with data

from Braslow and Visconti (1948). Granville used atransformationto convert this information to a bodyof-rotationproblem. The data were also correlatedwith turbulence level in the free stream as shown infigure 22. Extrapolationof the criteria woulda two-dimensionalairfoil dominated by TS(Holmes et al. 1983), whereby the existingincluded this form of transition. However,

Theory

method---correlatesthe predicted disturbancegrowthwith m_:asured transition locations. Although limitedto empirical correlationsof availableexperimentaldata, it:s the main tool in use through the 1990s.

work fortransitiondatabasewhen the

designconfigurationbegins to significantlydifferfrom the existing database, this transition predictioncriteria would likely fail.

Linear stability theory represents the current state

of the art for transition location prediction for threedimensional18

subsonic,

transonic,

and

supersonic

tions to predict the location

to turbulent flow.

flows. To begin a transition prediction analysis, the

steady, laminar mean flow must first be obtained(either by Navier-Stokessolutions or by boundarylayerequations).Thenthethree-dimensionalboundary-layerstability equations(Orr-Sommerfeldand Squire ordinary differential equations) are solvedfor the amplificationrate at each point along the surface, based on the assumptionof small-amplitudedisturbances.

are derived from solutions of a coupled Euler and

boundary-layerequationsolver. Harris,Iyer, andRadwan (1987) and Iyer (1990, 1993, 1995) presentedapproaches for the Euler and boundary-layerequationsolver. Harris, Iyer, and Radwan (1987) demonstratethe accuracy of a fourth-order finite-differencemethodfor a Cessna aircraft fuselage forebody flow, flat-plateboundary-layerflow, flow around a cylinder on a flat

plate, a prolate spheroid, and flow on an NACA 0012

swept wing. In terms of computationalefficiency, theEuler and boundary-layerapproachfor obtainingaccurate mean flows will be the solution of choice for

geometries,turbulence modeling efforts, and in thedirect numerical simulation of the unsteady flow physics (Kleiser and Zang 1991). However, a transition-

most of theNavier-Stokes

predictionmethodology devised in the 1950s is considered state of the art and is being used by industryfor LFC-relateddesignthrough the 1990s. Thistransition-predictionmethodologytermedthe e Nmethod is semi-empiricaland relies on experimentaldata to determine the N-factor value at transition.

preliminarydesign stages;however,solvers can be used for LFC design. A

limiting factor for the Navier-Stokes

mean flows is thedemanding convergencerequired for the suitability ofthe results in the boundary-layerstability codes.To obtain the stability equations, begin with thefull incompressibleNavier-Stokesequations that are

To derive the stability equations, take the velocities fi,_,& and the pressure _ as solutions of theincompressible,unsteadyNavier-Stokesequations.The instantaneousvelocities and the pressure may beinto base and disturbance

{fi,_,_v}(x,y,z,t)

components

= {fi,_,w}(x,y,z)+ {u,v,w}(x,y,z,t)

p(x,y,z,t)

where

fi(x,y,z)

the base flow is given

from laminar

The disturbanceevolution and transition prediction tools require an accurate representationof themean flow (velocity profiles). Either the velocity profiles can be extracted from Navier-Stokessolutions or

Significant advances have been made in the understanding

of the fundamentalsof two- and threedimensional,unsteady, viscous boundary-layerflow

decomposed

of transition

3fi

3;v

O'-x+ _yy+ _zz = 0

as

O-7+ U

+V

+W

1 (O2u

(5)

(6)

O2___fi+O2fi1

(7)

+ Oy2 Oz:)

+ p(x,y,z,t)

by the velocities

u, v, w

(8)

and the pressure/_,

and the disturbance component isgiven by the velocities u,v,w and the pressure p. Inthe Cartesian coordinate system (x,y,z), x is alignedwith the chordwise direction, y is normal to the wall,and z corresponds to the spanwise direction. To illustrate the stability tools, the Cartesian coordinate system and incompressibleequations are used herein. Ingeneral, curvilinearor generalizedcoordinatesareused to solve the govern system of compressible equa-

3y + Re_._x2

_y2

_z 2)

1=-_z

19

+Re_,bx2

by2

3z 2)

(9)

A Reynoldsnumbercanbe definedas

Re = USv,

where U is the velocity, 8 is the characteristic

and v is the kinematic viscosity.

d2

length,

For hydrodynamiclinear stability theory, whichmakes use of the quasi-parallelflow assumption,fi(y)and _(y) are functions of distance from the wall onlyand v = 0. Substituting equations (5) into the NavierStokes equations,

the following

-t_

-13

212[J-iRe

d2fi

d2_,

tx--_-_dy

+(O_u+_w-t.0)

-Or.

__2

2dy

_, =

dZfi

linear system results:

(15)

dfi

___ 2 _ i2

iRe(otfi + _

+ [_2 _

- to)] d-_-

ay

/)u

av

3w

oq'--x+ _y + _zz = 0

Ou

- Ou

dfi

(10)

= /Re

- 3u

lIa2u

- a-'--x+ Ree_x2

a2__u+_2./+ ay2

standard

wall boundary

-Or

Ow

a2V

_ee[,_x2 +--Oy

dw

a--7+.Uxx+,,_ap

and the free-stream

(a2V

-Ow

are

= 0

(y = O)

(17)

_Or

ap+l3y

conditions

(11)

at+U_x+Waz

(16)

_)Z 2 )

d_;,_y,Ov

@)a

+rdyj

where _ is the wall normal vorticity and dn/dy n "is

the nth derivative in the wall normal direction. The

at+u_+v_+Wazap

(0_

conditions

are

_2V)2

d_;'Uyy'

(12)

t)Z2_

+w_a2.______W+ )2W1ajx2+ay2 az2)

_ _ 0

(y _,,_)

(18)

Either spatial or temporal stability analysis may be

performed,whereby the temporal analysis is lessexpensive and the spatial analysis is more physical. Inaddition to the Reynolds number, Mach number, and

-Ow

l(a2w

az+_t,

boundary

(13)

other parametersanalysis requires

that must be prescribed, a stability

that the mean flow and its first and

second all-normal derivatives be known very accurately. ,t, small deviation in the mean flow could cause

Accordingto the conventionalnormalmodeassumptionused to derive the Orr-SommerfeldandSquire equation, the eigensolutionstake the form

significant changes in the second derivative and contaminate the stability calculation. Once the mean flowis obtained,

a stability

problem

has to determine

six

{_r,O_i,_r,_i,tor,_i},whichare thestream_ ise wave number, streamwise (spatial) growthrate, spanwise wave number and growth rate, wavefrequency, and temporal growth rate. For the temporalformula :ion, t_ and 13are real numbers and to is a comunknowns:

{u,v,w,p}

= {fi,_,,_,,/3I(y)

exp[i(otx+

_z-tot)](14)

where

i = ,f2-i, ct and

13 are

the

nondimensional

wave numbers (proportional

to wavelengths)in thestreamwiseand spanwise directions,to is the fre-

plex nm nber that is determined through an eigenvalue

solver. For the spatial approach, t_ and 13are complex,and to is the wave frequency.

quency, and {fi, _, _,,_ } describe the velocity profile.

Substitutingequation (14) into the linear equations(eqs. (10) to (13)), the following Orr-SommerfeldandSquire equations may be obtained:

Because the spatial formulation is more representative of the real boundary-layer

instability physicsand the temporal-to-spatialconversion is only valid on2O

the neutralcurve,the remainingtransitionprediction

methodologiesaredescribedvia thespatialapproach.However,thetemporalapproachwasintroducedfirstby SrokowskiandOrszag(1977)in theSALLYcodeandlaterby Malik (1982)in the COSALcode.TheCOSALcodeincludedtheeffectof compressibilityinthe equations.For the spatial approachin threedimensionalflows, the frequency (O r is fixed,(Oi = 0,and{_r,O_i,_r,13i}are parametersto bedetermined. Although an eigenvalue analysis will provide two of these values, the main issue with the application of the e N methodologyto three-dimensional

The

parameters

is assumed

available

Reynoldsnumbers,numbers.

to determine

amplitudeis tracked

fixed

wave

for three-dimensional

_}Ot i--

angle,

and

fixed

flows.

(20)

[3rThe group velocity

the

angle

_g

is

given by _O_r/_13r or

_g = tan -l(bo_r/_r

two remaining free parameters, the N-factor correlation with experiments

could be carried out. By integratingfromthe neutralpointwitharbitrarydisturbancedisturbance

point,

Strictly valid only in parallel flows, the saddle

point method suggests that the derivative of otx + _zwith respect to 13 equals zero. As noted by Nayfeh(1980) and Cebeci and Stewartson(1980), carryingout this derivative implies that dogd_J must be real or

The Reynoldsnumbers nondimensionallyrepresentthe spatial chordwise location on a wing (for exampie). The boundary between regions of amplification(unstable) and decay (stable) is termed the neutralcurve (location where disturbances neither amplify nordecay).If a method

saddle

spanwise wavelengthmethods are three approacheswhich have been devised to determine the two free

likely or expected for similar flow situations can be

inferred. The resulting N-factor is correlated with thelocation of transition for a variety of experimental data(sketched in fig. 24). This informationis then used indeterminingthe laminar flow extent (crucial to LFCdesign). Hence, this methodologyis critically dependent on the value of the experimentaldatabases andthe translation of the N-factor value to a new design.

(21)

The final condition to close the problem requires that

the growth rate be maximized along the group velocitytrajectory. Then the N-factor (or integrated growth)would be

A 0, the amplificationof theuntil the maximum amplitude

A 1 is reached at which a decay ensues. Being a linear

method, the amplitudes A 0 and A 1 are never reallyused; rather, the N-factor relation of interest is defined

N =

T ds

as

(22)

whereN=

A, =ln_o

T ds

(19)

where s o is the point at which the disturbance

firstbegins to grow, s I is the point at which transition iscorrelated, and Y is the characteristic growth rate of thedisturbance. Figure 24 illustrates the amplification anddecay of four disturbances(wave-number-frequencycombinations)leading to four N-values. The envelopeof all individual N-values leads to the N-factor curve.By correlatingthis N-factorcases, the amplification factor

--lOCi-

[_i

a13r)

y=

1+t713,)sO is the location where the growth rate g is zero, ands I is the distance along the tangent of the group velocity direction.

with many transition

for which transition is21

Forthenextmethoddeveloped by Arnal, Casalis,

and Juillen (1990), the fixed wave angle approach sets_i = 0 and the N-factorswave orientation or

N =

are computed

-o_ ds

fffl

In a discussion of the application of linear stability

theory and e N method in LFC, Malik (1987) describesthe methodologyfor both incompressibleand compressible flows and presents a variety of test cases. In

with a fixed

situations where transition occurs near the leading

edge of wings, the N-factors can be quite large compared _ith the range N = 9 to 11 applicable for transition in the latter portion of a wing. Malik makes an

important contribution to this understanding

Many calculationshave to be carried out over therange of wave angles to determine the highest valueof N.

The last method,

the fixed

spanwise

However, for transition near the leading edge of a

wing, the stabilizing effects of curvature are significant and must be included to achieve N-factors of 9to 11. The rest of this subsection documents samplesof the extended use of the N-factor method for predicting laminar flow extent.

wavelength

approach, proposed by Mack (1989) sets _i = 0 and

_lr is held fixed over the N-factor calculation, computed by equation (23). Many calculations have to be

Schrauf, Bieler, and Thiede (1992) indicate that

transition prediction is a key problem of laminar flowtechnology. They present a description of the N-factorcode developed and used at Deutsche Airbus, docu-

carried out over the range of 13r to determine the highest value of N. It is not clear what the significance ofholding

A major obstacle in validating or calibrating current and future transition prediction tools results frominsufficientinformation in wind tunnel and flight testdatabases. For example, Rozendaal (1986) correlatedN-factor tools for TS and CF disturbances on a flighttest database for the Cessna Citation III business jet.The database consisted of transition locations mea-

Am, rag others, Vijgen et al. (1986) used N-factor

linear stability theory to look at the influence of compressibility on disturbanceamplification.They compared TS-disturbancegrowth for incompressibleflowover a NLF fuselage with the compressibleformulation. They noted that compressibilityis a stabilizinginfluenc_ on the disturbances (1 st mode). For the NLFand LF,2, an increase in Mach number (enhancedcompressibility)is stabilizing to all instabilities forsubsoniv to low supersonic flow.

sured with hot-film devices for points that varied from

5 to 35 percent chord on both upper and lower wingsurfaces for Much numbers ranging from 0.3 to 0.8and altitudes ranging from 10000 to 43 000 ft. Theresults showed that CF and TS disturbances may interact and that CF disturbancesprobably dominated. CFN-factors were scattered around 5 and TS N-factorsvaried from 0 to 8. The stability analysis showed norelationshipbetween Mach number and disturbanceamplificationat transition.Rozendaal (1986) noted

Nayfeh (1987) used the method of multiple scales

to acco_ nt for the growth of the boundary layer (nonparalleleffects).The nonparallelresultsshowedincreased growth rates comparedwith the parallelflow assumption. These results indicate that nonparallel flow effects are destabilizingto the instabilities.Singer, Choudhari, and Li (1995) attempted to quantify the _fffect of nonparallelismon the growth of stationary :rossflowdisturbancesin three-dimensional

that the quality of the results was suspect because no

information on surface quality existed, an unresolvedshift in the pressure data occurred, and an inadequatedensity of transition sensors on the upper wing surfacewas used. Furthermore, the impact of the engine placement relative to the wing could be added as a potentialcontributing factor. The Rozendaal analysis reinforcedthat the N-factor method is reliant on good experimental data.

boundaE layers by using the multiple scales analysis.

The results indicate that multiple scales can accuratelyrepresent the nonparallel effects when nonparallelism

22

is weak;however,asthenonparalleleffectsincrease,multiplescalesresultsdiminishin accuracy.

describing the shape function

odology for predicting transition location. They note

that the main features lacking in the methodology arethe inability to account for the ingestion and characterization of the instabilities entering the boundary layer(the receptivity problem). In section 3.5.6, the issue ofpredicting boundary-layerreceptivity is discussed, butfirst, advance transition prediction methodologiesarepresented in sections 3.5.4 and 3.5.5.3.5.4. Parabolized

Stability

Equations

[n]

result. These equations

dx + [N]

Because

the fast variations

number,

the second

= f

(25)

of the streamwise

derivatives

are negligible. By the proper choice of O_n, m, this system can be solved by marchingin x. For smallamplitude disturbances,f = 0, whereas for finiteamplitude disturbances,f in physical space is simplythe nonlinear terms of the Navier-Stokesequations or

TheoryF = (u. V)u

Because the N-factor methodology based on linear

stability theory has limitations, other methods must beconsidered that account for nonparallelism,curvatureeffects, and ultimatelynonlinearinteractions.Thefinal method considered relative to the evolution ofdisturbancesor method.

Joslin, Streett, and Chang (1992, 1993) and Pruett

and Chang (1995) have shown that the PSE solutionsagree with direct numerical simulation results for thecase of incompressiblefiat-plate boundary-layertran-

N-factor method, which assumes a parallel mean flow,

the PSE method enables disturbance-evolutioncomputations in a growing boundary-layermean flow. Asfirst suggestedby Herbert(1991) and Bertolotti(1991), PSE theory assumes that the dependenceofthe convective disturbanceson downstream development events is negligible and that no rapid streamwisevariations occur in the wavelength,growth rate, andmean velocity profile and disturbanceprofiles. At

sition and for

respectively.

Zz n=-N z

I_m, n

dx+ m_z-noM)3

the total numbers of modes kept

Fourier series. The convective direc-

N t are

tion, or streamwise

direction,

has decomposition

a cone,

Finally, theoreticaland computationaltools arebeing developed to predict the rich variety of instabilities which could be growing along the attachment lineof a swept wing. Lin and Malik (1994, 1995, 1996)describe a two-dimensionaleigenvalue method which

(24)

where N z andin the truncated

on

by Joslin and Streett 1994 and Joslin 1995), whereas

the linear N-factor type results suggest that the disturbances continue to grow. Hence, the linear predictionsinadequately predict the behavior of the disturbances.

*m'n(X'Y)

X exp[i(J_x

transition

with the experiments

in that the stationarydisturbances reach a saturation state (confirmed with DNS

very much dependent upon the nature and spectrum of

the disturbanceenvironment,the signaturesin theboundary layer of these disturbancesand their excitation of the normal modes ("receptivity"),and finallythe linear and nonlinear amplificationof the growingmodes."

to Turbulence

In this subsection, a relatively new concept is outlined which involves coupling transitionpredictionmethodologywith a two-equationturbulencemodelapproach.transition

Warren and Hassan (1997a, 1997b) pose the

tem of equationsinvolving the kinetic energy andenstrophy. The exact governing equationsprovide alink between the laminar boundary-layerflow instabilities, the nonlinear transitional flow state, and the fullyturbulent flow fluctuations.If the breakdown is initiated by a disturbancewith a frequency reminiscent ofthe dominate growing instability, the simulations areinitiated. The influence of free-stream turbulence andsurface roughnesson the transitionlocationwasaccounted for by a relationshipbetween turbulencelevel and roughnessheight with initial amplitude ofthe disturbance.The initial comparisonswith flatplate, swept fiat-plate, and infinite swept-wingwindtunnelexperimentssuggesta goodcorrelationbetween the computationsand experiments for a variety of free-streamturbulencelevels and surfaceconditions.Approachesrelating flow instability andtransition and turbulence modeling show promise forfuturecomputationsof LFC-relatedaerodynamicconfigurations.3.5.6. ReceptivitymThe

Ingestion

Leeheyand Shapiro(1980),KachanovandTararykin (1990), Saric, Hoos, and Radeztsky (1991),and V_iegel and Wlezien(1993) have conductedreceptivity experiments;Kerschen (1987), Tadjfar andBodonyi(1992),Fedorovand Khokhlov(1993),Choudhari and Streett (1994), Choudhari (1994), andCrouch (1994) have conducted theoreticalstudies ofreceptivity to extend the knowledge base and capability for predictingthe receptivityprocess. Acousticnoise, turbulence, and vorticity are free-streaminfluences end couple with single and distributedroughness, s:eps and gaps, surface waviness,and otherthingsto producedisturbancesin the viscousboundary-layerflow which are relevant to NLF andLFC applications.These ingestion mechanismsarereferred to as "natural receptivity"; however, there areforced md natural categories of receptivity.Becausethe dominant instabilities in a boundary-layerflow areof a sh_rt scale, the receptivity initiation must inputenergy into the short-scale spectrum for the most efficient excitation of disturbances.As Kerschen (1989)pointed out, forced receptivity usually involves theintentional generation of instability waves by supplying energy to the flow at finite and selected wavelengths and frequencies that match the boundary-layerdisturbl,nce components. Examples of forced receptivity inchtde unsteady wall suction and blowing or heating and cooling (used for active flow control).

of Disturbances

Morkovin (1969) is usually given the credit for

coining the process called receptivity. Receptivityisthe process by which free-stream turbulence perturbsthe boundary layer by free-stream disturbancesoriginating at the edge of the boundary layer. Althoughbelieved by many to be a significant piece of the transition process, only brief mention is given to receptivity in this review. The rationale for this brief mentionlies with the fact that receptivity has not been an activepart in the history of LFC. However, receptivity willinevitably play an important role in the future of NLFand LFC technologies.Let us quote Reshotko (1984) for a description oftransition and the role of receptivity. "In an environment where initial disturbancelevels are small, thetransitionReynolds number of a boundary layer is

boundary-layerequations. In this situation, the tripledeck asymptotic approximationto the Navier-Stokesequations is used. The triple deck produces an interactive relationshipbetween the pressure and the displacementthicknessdueto matchingof therequirementsbetween the three decks. The middledeck or main deck responds inviscidly to the shortscale wall discontinuities.The viscous layer (lowerdeck) between the main deck and the surface is

where co is the frequency of the disturbance which one

desires to initiate, f(x) is the shape of the suction andblowing distribution (generally a sine or cosine bubbleshape), and v is the resulting wall-normalvelocitycomponent at the wall. Similar techniques can be usedfor unsteady thermal forcing and to excite disturbances in a wind tunnel experiment.

required to ensure that a no-slip boundary condition is

enforced at the wall. Finally, the rapid change in dis-

Natural receptivityis more complicatedin thatfree-stream acoustic, turbulence, and vorticity are ofmuch longer wavelengthsthan the boundary-layerdisturbance. Complicatingthe matter, the free-streamdisturbancein nature has a well-definedpropagationspeed and energyconcentratedat specific wavelengths. Hence, the free-streamdisturbancehas noenergyin wavelengthsthat correspondto theboundary-layerdisturbance.So a mechanismmusteffectively (and efficiently) be able to transfer energyfrom the long-wavelengthrange to the short wavelengths.Mechanismsto accomplishthis transferinclude the leading edge (of a plate and wing) and surface discontinuities(e.g., bugs, surface roughness,rivets).To determine (or describe)scaleconversion,Goldstein

placementthicknessat the surfacediscontinuityinduces a correction to the outer potential flow. Thiscorrection takes place in the upper deck. The meanflow gradients due to the discontinuityserve as forcingtermsfor the disturbanceequations.Therefore,although much understandingabout receptivityhasbeen gained over the past few years, significantresearch must be conducted,especially in the threedimensionaleffects and in supersonic flows, beforethe tools become widely used as design tools. Again,receptivity is included in this LFC review because itwill inherently play a role in future transition prediction for NLF and LFC design tools.

3.5.7. Optimize

this process of length

(1983,1985)and

Pertaining to the determination

of what "optimal"suction distributions should be used on LFC systems,

Goldstein, Leib, and Cowley (1987) showed that the

primary means of conversion was through nonparallelmean flow effects. Hence the two cases where nonparallel effects are strongestare (1) regions of rapidboundary-layergrowth as at the leading edge wherethe boundary layer is thin and rapidly growing and(2) downstreamat a surface discontinuitysuch as abump on the wall.

Nelson and Rioual (1994) posed a determination

bymeans of minimizingthe power requirementstoachieve transition at a specified location, by applyingsuction through a sequence of controllablepanels.Their paper had the problem formulated as a nonlinearconstrainedoptimizationproblem and focused moreon the stability of the algorithm than on the fluidsmechanics of the LFC system. In a comparable study,Hackenberg,Tutty, and Nelson (1994) showed convergence optimization of 2 or 4 panels is less than 10iterations for the problem of transition on a flat plate.

To determine the receptivity of the boundary layer

in the leading-edgeregion of a particular geometry tofree-streamdisturbances,solutions of the linearizedunsteadysolutions

Linear Design for LFC

boundary-layerequations are required. Thesematch downstream with the Orr-Sommerfeld

equation,which governsthe linear instabilityandserves to provide a means for determiningthe amplitude of the viscous boundary-layerdisturbance.

to predict the suction distribution under

the constraintof fixed mass flow (fixed energyrequirement). The beginning of the suction region wasimposed upstream of the neutral point and the end ofthe suction was prescribed downstreamof the transition point. For simplicity,the mean flow was

Th, applicationof thermal control for LFC aircraft is in an infancy stage comparedwith suctionLFC. I;sues relating to the thermal surface are unresolved as of this publication. One of these potentialissues involves the possibility of surface waves beinggenerated through the use of strips of thermal control.Whether such an applicationwould generate wavesintolerable to laminar flow has not been studied yet.

determinedby solving incompressibleboundary-layerequations. Although optimal suction is demonstratedfor TS wave control in a flat-plate boundary-layerflow (Blasius), the resulting suction distributionsfortravelingand stationarycrossflowdisturbancesinswept Hiemenzflow are quite relevant to HLFCimplementedon a swept wing at low speed. Interestingly, the region of maximum suction occurred verynear the location of the onset of disturbance amplification and progressively decreased through the region ofdisturbancegrowth. In addition, Balakumar and Hallconcluded that over an order of magnitude more suction is required to control crossflow disturbances compared with that required to control TS disturbances.

3.5.9. AdvancedTolerances

Prediction

of Manufacturing

Innovativetools have been developed to predictthe impact of manufacturingtolerances on the extentof laminar flow; however, very little validationofthese tools has been documented.As Masad (1996a,1996b) shows, interacting boundary-layer(IBL) theory, which accounts for the viscous-inviscidinteraction, can be coupled with either linear stability theoryor PSE theory to parameterizethe allowable dimensions of steps, gaps, rivets, and other things, which canbe used and not impact the laminar flow.

Stock (1990) posed an interesting way of viewing

boundary-layerinstability with suction. The problemwas transformedfrom the problem of a boundarylayer flow with a pressure gradient and suction to theproblem of an equivalentpressure gradient withoutsuction. The equivalence is imposed based on an identical form parameter, or shape factor H. Using integraland finite-differencemethods, the stability results forthe case with and without suction were shown to be in

Nayfeh,Ragab, and AI-Maaitah(1987,1988)looked at the issue of manufacturingtolerances by performing a study of boundary-layerinstability aroundhumps and dips. Interactingboundary-layertheorywas ust:d to account for the viscous-inviscidinteraction associated with potential separation bubbles, andthe amplificationof disturbancesin the presence ofhumps with various height-to-widthratios and at various locationswas studied. The results suggest thatN= 9 correlates well with the transition location. In

agreement.

3.5.8. Thermal LFC

additior:, the size of the separation bubble is influenced by the height-to-width

ratio and Reynolds number, anti the disturbance instability is affected by theheight-to-widthratio and the location of the imperfection from the leading edge of the plate and branch I ofthe neutral curve.

As early as the 1950s, the thermal concept was

recognized as a potential means for boundary-layerstabilization.Dunn and Lin (1953) realized and demonstrated that mild surface cooling was able to stabilize viscous boundary-layerinstabilities which wouldotherwise amplify and lead to transition. In fact, thecalculationsshowed that 2D disturbancescould be

4. Laminar

Flow

Control

Aircraft

Operations

completely stabilized at Mach number of 1.6 for the

ratio of wall to free-streamtemperatureof 1.073,which implies a small amount of cooling.

The operationalincludirgcontrollinginsects, is paramount

A more recent study by Boeing (Parihk and Nagel

1990), showed that with stability theory cooling canbe stabilizing to both TS and CF disturbanceswithapplication to supersonic LFC transports.

aircraft. Both ice and insects generate

roughnessinduced prematureloss of laminar flow. Althoughanti-icingsystems have been operationalfor manyyears on the leading edge of wings and on nacelles,26

only limited research

results for realisticinsectprevention systems are available. This section focusesprimarily on the issue of insect accumulationand prevention; brief discussions on aircraft icing research,

nighttime collections,in October followed

the impact of atmospheric

particulateson laminarflow, and boundary-layercontrol for high lift willfollow. Finally, a discussion of operational maintenance of laminar flow closes this section.

decreasingrapidlywith increasedaltitude.Glick(1939) also noted that temperature was one of the mostimportant meteorologicalfactors in the control anddistribution of insects. He showed that the maximum

4.1. Insect

period indicated that the largest density of insects was

measured at low altitudes, with the number of insects

densities were measured

at temperaturesof 75 to80F. Finally, Glick (1939) noted that the insects andmites captured at high altitude (and one spider at15000 ft) were very small and completelyat themercy of the air currents. The size, weight, and buoyancy of the insects contributed directly to the height towhich the air currents carried it and hence to the pres-

ence of insects at high altitudes.

Hardy and Milne (1938) reported on the distribution of insects with altitudes from 150 to 2000 ft. The

this issue follows.

measurements

populationdensity qualitativelyagreed with Glick(1939) in that the largest density was at low altitude.

UnitedStates (lowerMississippivalley)and atTlahualilo, Durango, Mexico. Much of the test area isswamp country,encompassinghundredsof smalllakes, bayous, rivers, and great forests. The project ofcollectinginsect data was conductedfrom August1926 to October 1931 and the results are reported by

Although all insects were affected somewhat differently by the weather conditions, high temperature andlow humidity were determined to be more favorable toaerial drift than the reverseconditions.Freeman(1945), under the direction of Hardy, expanded on theearly kite-flownstudy and found that the greatestnumbers and varieties of insects occurred in May,

Glick (1939) in a Department of Agriculture Technical

Bulletin. The investigationis of importanceto LFC(and aircraft in general) and documented the numbersand kinds of insects, spiders, and mites with atmo-

June, and September.

Although the information inthese studies were significant for the NLF technology,the primary goals of the studies focused on characterizing the insect families and the motion of agricultural"pests" from one location to another.

armysometrapswere

placed between the biplane wings. A protective cover

was used to control the duration and altitude of exposure to the screens. All measurementswere made with10-min exposures

were made with traps and nets carried

into the sky by kites in England. Their study conducted from 1932 to 1935 resulted in 839 insects captured in 124.5 hr of flight. Of interest here is that the

On August 10, 1926, the first known attempt to

use an airplane in collecting insects was made underthe direction of E. P. Felt at Tallulah, Louisiana, in the

spheric conditions and altitude. DeHaviland H1

biplanes were used for the study and covered150 000 miles. For the measure of insect density,of 1 ft 2 embedded with fine-mesh copper screens

the greatest numbers were taken

by May. Results over the 5-yr

Incidentally,in theflighttestingoftheHurricane II reported by Plascott et al. (1946), no fliesor insect debris was observed in this NLF flight test.However, the drag measurementsfrom previous flighttests where flies and insects were picked up indicatedan increase in the drag due to insect debris. Hence, thefull advantagesof laminar flow and the subsequentlow drag would require some method to prevent theinsects from adhering to the surface.

at known speeds.

Although the altitudes ranged from 20 to 16 000 ft,

the systematic studies were conducted at 200, 1000,2000, 3000, and 5000 ft for daytime collectionsand1000, 2000, 3000, and 5000 ft for nighttime collections. Over all altitudes, Glick (1939) reported that thegreatest number of insects was taken in May, withNovemberand Septemberfollowing.The fewestinsects were taken in January and December.For

the test section

with paper was the sim-

early in the morning). This

increase with temperature

discussed earlier. To avoidthe insect problem, a sheet of paper covered 0 to30 percent chord on the upper and lower surface of thetest section. After the aircraft takeoff and climb to sufficient

altitude,

the pilot could jettison

the paper

by

pulling a string attached to the paper and retrieving the

paper inside the cockpit through a piece of pitottubing.

To avoid

insect

contamination

for the Vampire

porous-suctionflight tests reported by Head, Johnson,and Coxon (1955), the test-sectionsleeve was protected during takeoff and climb by a strip of tracingpaper that covered from the leading edge to about10 percent chord and was fixed to the surface withadhesive tape. Takeoff was delayed until 100 knotshad been reached. This speed was maintained duringtakeoff and climb, and at "sufficientaltitude," thetracing paper was jettisoned90 knots.

by reducing

the speed to

To avoid insect contamination

for the F-94A flighttests of a slot-suctionLFC experimentreported byGroth et al. (1957), the first 30 percent of the upperand lower surface of the test section on the wing wasprotected with a cover of blotting paper taped to thewing. This paper remained attached through takeoffand climb, then the plane was deceleratedto removethe covering. Without this covering, turbulent wedgeswere generated from the insect remains. However, fulllaminar flow could be regained by climbing to higheraltitudes (25 000 ft). This regaining of laminar flow isunderstood to be a unit Reynolds number effect. Forconstant Mach number, a climb in altitude decreasesthe unit Reynoldsnumberand, as discussedinsection 2, a lower unit Reynolds number flow is moretolerant to a roughness (insect impact) of given size.

plest (or least mechanical)

anti-insectdevice. Thisdevice was successfullyused in the major laminarflight tests, including Gray and Davies (1952) with theKing Cobra flight test; Head, Johnson, and Coxon(1955) with the Vampire porous-suctionflight test;Groth et al. (1957) with the F-94 slot-suction flighttest; and RunyanNLF flight test.

flight tests were conducted

observationis consistent

For the Boeing 757 NLF flight tests (Runyan et al.

1987), the glove was protected from insect strikes during takeoff and climb by using a paper covering untilthe airplane reached 5000 ft at which time the paperwas pulled into the cabin via a nylon cord. On flightsnot using the protective covering, loss of laminar flowwas observed during the flight and evidence of insectaccumulationnear the attachment line was measured

with the Boeing 757

Gray and Davies (1952) reported on King Cobra

flight tests at the Royal Aircraft EstablishmentinEngland.As the Spring days became warmer, theinsect contaminationproblem increased (even if the

after landing.29

4.1.2. Scrapers

describ._d

Wires and felt pads have each been tested with

some success in wind tunnel experiments,the latterworking for painted surfaces. The problem of dragpenalty due to the device was not evaluated; however,the device must either be contained in the skin of theaircraft during cruise flight or be jettisonedunreasonable

drag penalty (Coleman

in section

6.6. The results indicated

that it

was po,_sible to protect the upper surface but it was not

possible to protect the lower surface. However, theplate device caused considerabledrag and a pitchingmoment. The retracted reflector could introduce significan! ridges. The Krueger flap serves to both protectthe surface from insect strikes and improve lift.

to avoid an

1961).

4.1.3. Deflectors

4.1.4. Fluidic

Cover

Coleman

(1952)

discussed

wind tunnel tests that

employed the application of glycerine, glycerine and

gelatine, and soap and methanolto wing sections.These solutions would be wiped away as the aircraftreaches sufficient speed to cause the shear to removethe fluid (and insects). Although these solutions wereshown to decrease the accumulationof insects on the

Deflectorsconsist of a surface (or plate) thatforms a nose flap which protects the leading edge ofthe wing from insects and absorbs the insect impacts.Tamigniaux,Stark, and Brune (1987) discussedawind tunnel experiment to test the effectivenessof theKrueger high-lift device used as a shield againstinsects (although the insects used were larger relativeto the model size than would be encountered in flight).Note, figure 2 shows a leading-edgeKrueger device,which would be retractedafter takeoff and climb,

test article, complete elimination of the insects was not

possible. Continuousspraying of the solution wasshown Io be effective and required a penalty of 0.2 to0.5 of the TOGW of the aircraft.

leaving a clean leading edge for cruise. The 2-ft model

consisted of a slotted-leading-edgeKrueger flap on awing section. The insects were injected into the windtunnel at a free-stream velocity of 4 ft/sec upstream ofthe wing leading edge. Without insects, the Kruegerflap was varied for 37 different positions, optimizingfor maximum high-lift characteristics.The optimalposition was a 45 deflection and the optimal gap andtrailing-edge gap were both 2 percent of the airfoil reference chord. The results showed that lighter insectsimpacted farther aft of the stagnation line than heavierinsects;this indicatesthat heavierinsectshave

4.1.5. Thermal Cover

straighter trajectories than lighter insects. A particle

trajectorycode was developed for two-dimensionalmultielementairfoils; the calculatedresults were in

takeoff and climb by the conventional

de-icing systems. Tie application of the ice layer to the aircraft,potentially damaging effects of large ice pieces breaking away from the wing, the required thickness of icerequired to prevent insect contamination,the minimum time to remove the ice layer, and the associatedperformancepenalty during takeoff are issues thatmust be addressed. Coleman (1952) discussed some

Under the concept of thermal covers, flammable

covers which could be electrically ignited can be rendered out of possible solutions because of safety (andpollutioa)concerns.Heating(rather cooking)theinsects antil they are consumedhas been suggested,but the high temperaturesrequired would be undesirable to the wing structure. Imposing a layer of ice onthe structure has been suggested and such a conceptwould be ideal in terms of preventing insect accumulation. Tl-is layer of ice would then be removed after

good agreement with the experiment. Insects impacting at an angle less than 7 left negligible body remnants on the wing upper surface to trip the laminarboundary layer. The Krueger concept has been demonstrated to be effective in flight on Jetstar LEVI" aircraft(Powell 1987); however, incorporatingan anti-icingsystem into the Krueger device remains an issue.

wind tulmel tests addressing

some of these issues.

4.1.6. RelaminarizationThis concept has been developed into the modemday Krueger flap and demonstratedon the Jetstarflight test (Maddalon and Braslow 1990) described in

Coleman

(1961)

noted

that

relaminarization

through the use of suction slots was investigated

byCumming, Gregory, and Walker (1953). The results of

section 6.3. Also, this concept was successful for the

Boeing 757 HLFC flight experiment (Collier 1993) as

their wind tunnel experiment

30

indicated

that the pump

dragincreasedbecauseof

of Coleman,

the suction approximately

balancing the profile drag due to the insect-roughenedsurface; hence, no apparent performancegain wasrealized with the suction slot.

1952). Once insects have accumulated

dry surfaces, they could not be removed

water and detergent spray.

in flight with

In the Croom and Holmes (1985) flight

ment, three different fluids were considered4.1.7. Liquid Discharge

glycolmethylether, and (3) monoethyleneglycol(MEG) and water. The fluid was discharged througheither slots or perforated holes, where the holes hada diameterof 0.0025 in. and were spaced about0.0205 in. apart. The TKS anti-icing system served asthe method for the current test, partially because thesystem has already been certified for several aircraft.The left wing which had no insect protection was usedas the baseline. The tests showed that the insect-

to estimate the insect accumulation

problem at LosAngeles,Sacramento,and San Franciscoairportsunder normal airline-typeoperations.Insects wereaccumulated on 13 of the 15 flights and caused premature transition. The initial flights confirmed that insectaccumulationandresultingprematuretransitionrequired an anti-insectaccumulationsystem. At the

protectionsystem should be activated before insectimpact. The ratio of water to MEG in the fluid systemand the flow rate played significant roles in the effectiveness of the insect protectionsystem. The MEG/water solution of 20/80 percent was very ineffective inreducing the number of insect strikes. Approximately10-percent fewer strikes were realized by using thissolution.However, with 80/20 percent solution,a75-percent(or greater) reductionin the number ofrecorded insect strikes was realized. As the flow rate

trailing edge of the flaps, boundary-layer

probesrecorded the state of the boundary layer. Next, fivespanwise segmentsof the leading-edgeflap weretreated with (a) an aluminum alloy untreated surface,(b) a spray-on DuPont Teflon coating, (c) DuPontTeflonpressure-sensitivetape, (d) organosiliconehydrophobiccoating, and (e) random rain repellentcoating. Flights were then conducted from many airports in the United States ranging from California toTexas to Florida. Insects were encounteredon all

wasincreased,thetotaldecreased. Croom and Holmes

insect(1985)

accumulationnoted that only

a 3-in. perforated region on the panel and a flow rate

of 0.16 to 0.33 gal/min were required to achieve a 68to 82-percent reduction in the insect accumulation.BulgubureandArnal(1992)andCourty,Bulgubure, and Arnal (1993) reported the use of aTKS insect avoidance system for the HLFC flight testsusing a Falcon 50 test aircraft. Monopropyleneglycol(MPG) was the fluid chosen for use in this system.

flights and the coatings were insufficient

to removethe insect contaminationinterruptinglaminar flow.The insect accumulationon super-slickTeflon surfaces and hydrophobiccoatings was comparedwithstandard reference aluminum. The flight test resultsshowed that none of the surfaces tested showed anysignificant advantagesin alleviating the insect contamination. Five types of flight tests were conductedwith the spray insect-avoidancesystem: (1) no spray,(2) water-detergentsprayafter all low passes,(3) large-dropletwater detergentspray after lowpasses, (4) continuous water spray during low passes,and (5) intermittent water-detergentspray during twopasses. The first test was used as the calibration or reference flight. The flight test with continuousspraywas most effectiveand no insect remainswerein the spray area (consistent

experifor the

purpose of both insect prevention and ice protection.

Thesolutionswere(1)monoethyleneglycol(Aeroshell07) and water solution,(2) propylene

Peterson and Fisher (1978) reported on insect contamination by using a Jetstar aircraft. The goals of theexperiment were investigating the extent of the insectproblem at large airports, determine whether insectaccumulationwould erode in cruise flight, test theability of the then new surface coatings to alleviate theinsect accumulationproblem, and test leading-edgesprays for anti-insect protection. In November1977,the Jetstar was flown on 15 takeoff and climb missions

observed

on

During low-altitudeflight tests over insect-infestedareas, the port (untreated)side of the aircraft had600 insects/m 2 impact the leading edge in the regionof interest, whereas on the starboard(treated) sidewith the MPG fluid, no insect contaminationwasnoted. Hence, the TKS system was very effective forinsect avoidance.4.1.8. Flexible

to perform ground-basedtesting. Additional effort wasplaced 9n accompanyingsimulationtools to predictthe accumulationand prevention.Refer to Britton1990, Perkins and Rieke 1993, and Bergrun 1995 fordiscussion of the icing issues; to Reinmann 1981 for abibliographyof ice-related research; and to Ranaudo,

Insect contaminationis usually limited to theleading-edge region from 0 to 30 percent chorddensity

of insects

falls

and Atmospheric

The National Advisory Committee

for Aeronautics (no_v the National Aeronautics and Space Administration)startedstudyingthe accumulationandpreventLon of ice on aircraft in 1928. An icing researchtunnel was built at the Lewis Research Center in 1944

4.2.1. Ice Accumulation

From these studies, we find that predicting and

The modern-dayKrueger flap can be used for insectpreventionand for increased lift during takeoff andlanding.

The accumulationof ice on the leading edge ofwings can significantly alter the geometry of the wingand cause drag penalty and performancedegradation(and in the worst case, safety can be affected). In addition, degradation of laminar flow can occur due to particulates in the atmosphere, most evident during cloudencounters.

if a test article covered with a coat-

ing designed to repel insects (similar to the concept by

Wortmann1963) would solve the insect-adhesionproblem for NLF and LFC applications.Subsequentflight tests with a NASA Learjet were carried outunder a cooperative agreement between NASA Lewisand Langley Research Centers and the General Electric Company.The results are not available for thispublication.

Of the anti-insect devices tested, paper coverings,

continuous liquid discharge, and deflectors have beendemon, trated in flight to prevent insect accumulation.Anti-ic ng systems such as TKS can be used to reducethe impact of insect accumulation.Solutions of MEGand water prevents insect accumulation(up to 82 percent) but is rather ineffective in removing insects fromthe surface after adhesion. Reduced insect accumulation occurs with increased solution fluid flow rates.

Reehorst,and Potapczuk1988 for a more recentreview of the NASA Aircraft Icing Research Program.AlthouGh much researchhas been performedforstandardconfigurations,little has been done forLFC-related aircraft.

extent in the Jetstar flight test program.

A cloudparticle spectrometer(Knollenberg probe) and a particle detector (charging patch) were used to measure thefree-stream particle environment. A degradation of theflow was observed during a cloud encounter coinciding with a charge-currentincrease on the instrumentation; however, full laminar flow was regained within afew seconds after the cloud encounter. Indicated by

noted the impact of

atmosphericparticles on achieving laminar flow during the flight test. Figure 26 shows a sketch estimatingthe LFC performance with ice particles in the air. Thefigure indicates that ice particles can influence laminarflow if the size and density of particles are sufficientlylarge. The flight results indicated that laminar flowwas lost as the size and density of particles increased.

Fisher and Fischer (1987) and shown in figure 27, the

Jetstar ice-encounterresults agreed with the Hallcriteria.

Hall (1964) set out to explain why the X-21 LFC

flight experiment lost laminar flow when the aircraftflew through visible clouds. The explanation began bylooking at the impact of the wake from a discrete particle on the otherwise laminar boundary layer; this sug-

Finally,

Anderson

and

Meyer

(1990)

showed

flight data for the F-14 NLF flight experiment

thatindicated turbulent bursts were measured during cloudencounters. The charge patch indicated the presence ofice particles during the loss of laminar flow while inthe clouds.

gests that local turbulent spots could be initiated in the

boundary layer, depending on the particle Reynoldsnumber and geometry. Next, the impact of surfaceroughnesswas reviewed, concluding that the roughness did not affect the boundary-layerstability belowsome critical roughness height or roughness Reynoldsnumber of 600 for spheres (3D roughness) and 200 for

Meifarthand Heinrich(1992) discussedissuesrelating to maintainingNLF and LFC in flight. Inagreement with the insect-contaminationissue at lowaltitudes, figure 28 suggests that atmospheric pollutionmay be an issue at high altitudes, even up to 10000 m.The uncertainty of the reliability of LFC systems operating in a polluted environmentcould be an additionalrisk to the implementationof the technology on a commercial transport; however, no degradation of the laminar flow extent was observed for the Jetstar LEFT test

cylindrical roughness (2D roughness). From the experiments, Hall concluded that the local boundary-layerReynolds number, pressure gradient, and free-streamturbulence had no effect on the critical roughness Reynolds number; however, an increase in Mach numberled to an increased critical Reynolds number. Fromthis review, Hall concluded that transition induced bythe wake of a particle was a local effect independent ofthe usual parameters (e.g., pressure gradient) influencing boundary-layertransition. To connect this impactof particles and roughness to the loss of laminar flowon the X-21 experiences, the particles in the cloudsmust be of sufficient size and density for sufficientduration to produce and sustain turbulence. Based onsparse data, the ice crystal size, density, and length ofexistence observed in the atmosphere correlated with

(see sectionpollution,

6.3) even though the Jetstar

dirt, and so forth at the various

4.3. Boundary-Layer

Control

encounteredairports.

for Takeoff

and

LandingAlthough boundary-layercontrol (BLC) is beyondthe scope of this review, a comment will be made here33

resulting pressure drop was 10 psf for supersonic LFC

and 20 to 40 psf for BLC. The BLC led to a dragimprov,_ment of about 10 percent over the optimizedflap configuration.Parikh et al. (1990) noted that amore definitive assessmentof performancebenefitsdue to BLC should be made through wind tunnel tests.4.4. OperationalFlow

be applied

more

Maintenance

of Laminar

The maintenanceand manufacturingof smoothsurfaces is a significantissue in achieving laminarflow, potentially creating an additional burden on theday-to-day operations of NLF and LFC aircraft.

Parikh et al. (1990) did a Euler computational

analysis of the BLC suction concept with applicationto a supersonic transport. An assessment of the impacton aerodynamicperformance with BLC was comparedwith the simple flap device without BLC. Boeing's 3Dinviscid flow code--PANAIR--wasused for a portion of the study. The Euler analysis was deemed sufficient for the study since previous studies have shownthat the inviscid analysis was capable of capturing thevortex formationand nonlinear evolutionon sharpleading-edgewings. The Euler analysis provided thepressure distributions,which were then used in a 3Dboundary-layeranalysis to determine the state of theviscous flow. The significanceof Reynolds numberscaling was an important factor drawn out in the analysis. At flight Reynolds numbers, the inboard portionof the wing indicatedattached flow. However,atlower Reynoldsnumbers (but same unit Reynoldsnumber), the flow separated on models which wereless than 1/4-scale. The calculations were repeated toinclude unit Reynolds number variations. The conclu-

Gray and Davies (1952) reported on the experiences gained at the RAE in England dealing with surface deteriorationissues. In the King Cobra flighttests, the test section of the wing was coated with twocoats of primer and one coat of filler, followed byadditional smoothing when deemed necessary. Over a6-month period, the surface deterioratedonly in theskin joints regions. The aircraft was exposedtoweather for about 200 hr and 50 flights entailing about40 hr. The rest of the time it was housed in a hangar.For different King Cobra aircraft, which was in theopen for about 2 years, the skin surface was chalky(dirty) _nd rivet and joints areas were the only areas ofthe wing that had any surface damage (cracking). Thesurface degradation results at the rivet-gap-jointareaswere consistent with those found by Plascott (1946)and Plascott et al. (1946) for the Hurricane II flighttest program. Gray and Davies (1952) noted that oncethe ground crews became habitually aware of the sensitivity .'equired for handling the wing surface for theHurricaae and King Cobra programs, protective coverings for the surface became unnecessary.

sion was that flow separation was only impacted by

chord Reynolds number effects. However,the unitReynoldsnumbercalculationsdid not take intoaccount the additional sensitivity of the flow to roughness (steps, gaps, joints). For the outboard portion ofthe wing, separation was encountered (when transitionwas assumed to occur at 5 percent chord). The effectof BLC and suction-regionextent were then studiedfor the separatedflow problem.The "optimized"results showed that for the four spanwise regions studied, a chordwise extent beginning at the suction peaklocation and covering1 percent chord was sufficientfor separationcontrol.The resultsshowedthat

In the description of a porous-suction

flight experiment on a Vampire aircraft, Head, Johnson, andCoxon q1955) noted an operational issue that must beaddresse.d when using powered suction systems. If thesuction pump were to fail, then outflow could causeprematureseparationat high lift coefficients.Thispotential problem could be alleviated with simple nonreturn valves to prevent outflow conditions.

plane. The tests were confined to the second slot of

(1992) had an in-depth

dis-

Some issues include the need for additional spare parts

and maintenance due to the suction system, uncertainties in the potential contaminationdue to pollution residue on the structural surface, and operational plan forsuction-systemfailure. The latter concern affects adecrease in range and increase in fuel burn as a resultof the unexpected turbulent drag.

Flow Control

Prior to

Oil Embargo

In this section,

LFC projects

are discussed

for the

time frame prior to the OPEC oil embargo. Each section has the configurationor modelinformation,project goals, and summarized

only after the suction was increased beyond 2.4 times

the normal value; (3) for the 0.015- and 0.03-in. plugs,normal suction produced turbulence and reducing the

5.1. B-18 Slot-Suction

level by 80 percent reestablished

laminar flow; (4) for0.2-, 0.5-, and 1.0-in. plugs, greater than normal suction values were limiting; and (5) the upper suctionlimit increasedwith increasingReynoldsnumber.Essentially, the slot blockage can cause a pair of adjacent vortices to combine and form a horseshoe vortexand lead to turbulence.

flight test experiment

were reported in an NACAWartimeReportbyZalovcik,Wetmore,andVon Doenhoff (1944). A test panel with nine spanwisesuction slots was mounted on the left wing (NACA35-215 airfoil) of a B-18 airplane (provided by theArmy Air Corps). The test panel shown in figure 29had a chord of 204 in. and a spanwise extent of 120 in.at the leading edge and tapered to 60 in. at the trailingedge. The nine original suction slots were spaced

Because the X-21A wings were built from many

panels spliced together on the wing, epoxy fills wererequired over the panel splices to meet the high unitReynoldsnumberstep and wavinesstolerances(Fowell and Antonatos,1965). However, the epoxyencounteredcracking and chipping under the wing

5 percent chord apart and were located from 20 to

60 percent chord. The eight additional slots were lateradded between each of the original slots. Suction wassupplied by an 85-hp Ford engine. Below each slot,the external flow was drawn through 0.25-in-diameter

loading and temperature

changes of flight. The bonding process proved to be the cause of the fill unreliability and the process was successfullychanged toachieve reliable tolerances.However,most of thetime was charged

laminar

cussion of issues relating to achieving and maintaining

NLF and LFC from the operations perspective. A flowchart of multidisciplinaryissues whichmust beaddressed prior to the use and reliance of laminar flowon aircraft performancewas presented. Issues whichwould cause an increase in DOC for aircraft and thosewhich would cause a decrease in DOC are connected.

OPEC

rized as (1) for the 0.007-in. plug, no turbulence was

observed for the range of normal to maximum suction;(2) for the 0.0115-in. plug, turbulencewas realized

maintenance

Meifarth and Heinrich

5. Laminar

were individually tested. All slots maintained the normal suction distribution,whereas the suction in theslots in chamber 5 was varied. The results are summa-

ground

and maintenanceof these joint areas. Furtherflow tests must carefully address this issue.

holes drilled in the wood panel spaced 0.75 in. apart.

The airflow was manually regulated by butterflyvalves located in the cabin. Static-pressureorifices

to the repair

35

locatedin theducts

or tubes were used to measure

the

sketch of the bronze porous sheet covering

a coreNACA64A010airfoilmodelperforatedwith1-in-diameterholes over the center of the model and

airflow through the slots. Numerous

coats of paint,filling, and sanding were employed to smooth the surface and to achieve an acceptablesurface-waviness

l-in. s;its at the leading and trailing

model. Suction airflowmeasurements

limit. Five-tube rakes were used to measure boundarylayer profiles, and two-tube rakes were used to measure the transition location.

through

an orifice

edges of thewere made

plate in the suction duct,

and suc-

tion was regulated by varying the blower speed and

plate orifice diameter. Boundary-layer measurementswere made on the upper surface to 83 percent chord.Laminar flow was observed to 83 percent chord forsuction up to a Reynoldsnumber of 8 x 10 6. An

The flight tests were conducted for chord Reynolds numbers between 21.7 106 and 30.8 x 106 withairspeeds from 147 to 216 mph. Uniformly increasing,level, and uniformly decreasing suction in the chordwise direction were applied. Laminar flow back to45 percent chord (pressure minimum point) was mainmined over the range of Reynolds number and lift

accompanyingtheoretical study suggested that, in theabsence: of roughness, full-chord laminar flow shouldbe expected to higher Reynolds numbers if the experimental suction distribution could be made uniform.

coefficient for suction mass flow Cq of 1.7 x 10-5 in

slot 1 and decreasing to almost zero suction in slot 5.If suction was further decreased in slot 5, reverse flowin that slot led to abrupt transition. Increasing the levelof suction had no additionalfavorable or adverse

In a follow-on test, Braslow et al. (1951) reported

the wind tunnel results of an experimentusing thesame model but with less porosity. Full-chord laminarflow v_as observedup to a Reynoldsnumberof24 x 106. The measured drag for the laminar flow control airfoil was roughly one third of the model without

effect on the transition point. However, for uniform

level or increasingsuction distributions,a criticalmaximum level of suction (Cq > 3.5 x 10 -5 in slot 1)led to turbulence regardless of the flight conditions.Finally, the results with 17 slots (2.5-percent-chordspacing of slots) were inconclusivebecause several

suction: however, the results could not be repeated

because the bronze skin buckled during testing.

small chordwiseof the panel.

5.2.2. UniversityTunnel Tests

5.2. LFC

cracks appeared

Wind Tunnel

near the leading edge

windLFC

5.2.1. Wind Tunnel Test With Porous Bronze Airfoil

Because Braslow, Visconti, and Burrows (1948)indicated that suction through a porous surface couldlead to performancegain, Braslow et al. (1951) conducted a LFC experimentinvolving a porous-suctionmodel in a low-turbulencewind tunnel. Using a modelwith a 3-ft chord and 3-ft span, experiments were carried out in the Langley Low-TurbulencePressure Tunnel (LTPT). The upper and lower surfaces of themodel were constructed from a single sheet of continuous bronze giving a single joint at the trailing edge.An estimate of the surface waviness indicated that_+0.003-in. variation occurred between the bronze surface and the inner aluminum

shell. Figure

Slot-Suction

Wind

Pfe:minger,Gross, and Bacon (1957) describedthe results of the LFC slot-suction experiments in theUniversity of Michigan5-Ft. by 7-Ft. Tunnel conducted in 1949 and 1950. Suction was applied through86 fine slots from 25 to 95 percent chord on a 30 swept 12-percent-thicksymmetric wing model. Totalpressure,static pressure,boundary-layercrossflow,and the transition location were measured during theexperin'ent.Measurementswere made at variousReynok_s numbers for model angles of attack of 0 and __1c. The suction for each test case was selected

Tests (1949-1963)

This section describes the early subsonic

tunnelexperimentswhichfocusedon thetechnology.

of Michigan

basedon theory.Full-chordlaminarflowwasobservedat an angle of attack of 0 at a chordReynok s number of 11.8 x 106. The measured minimum cr tical suction levels

were slightly smaller

than

theoreti_:al predictions;however, the measured dragclosely _natched the theoretical predictions. The suction lew:l on the 30 wing was slightly larger than a2D wing because crossfiow disturbances had to be stabilized. At an angle of attack of -1 , turbulent burstsoccurred for lower Reynolds numbers; this was correctly attributed to stronger crossflow.

30 shows a36

5.2.3. Douglas

Slot-Suction

was easily obtained up to a chord Reynolds number of

6.5 x 106 in the TDPT. Laminar flow was progres-

Wind Tunnel Test

sively lost with an additional increase in wind tunnel

speed. Hot-wire surveys behind each slot revealed thepresence of wild disturbancesbehind slot 6 (55 percent chord), which were most likely attributable to a

Smith (1953) presented

a review of LFC/BLCresearch at the Douglas Aircraft Company and notedthat the program began early in 1948. The studies suggested that as the Reynolds number increased the slotsmust become thinner and thinner; this caused doubt

0.003-in. step. Great care was then taken to remove all

discontinuitiesin the model. Additional tests showed

about the structural feasibility of the concept. Smith

conceived the idea of having several velocity discontinuities and regions of favorable velocity gradients forboundary-layerstabilization. However, such an airfoilmust not separate if suction power was lost. Thenature of the concept may cause shock formationateach jump; however, the suction would be sufficient toprevent separation.

that laminar flow was again lost, even though the flowwas theoreticallystable to TS disturbances.Theresults suggested that the flow was very sensitive tosurface roughness. Because of the surface-roughnessproblems, the test data were insufficient to make anyconclusionsaboutthesawtoothpressure-jumpdistributionconcept combinedwith slot suction forBLC/LFC.

To test the concept, a 2D airfoil (G00107) model

was installed in a Douglas wind tunnel. The wind tunnel could reach a maximum Reynolds number of4.25 x 106 and had a maximum fluctuating velocity of

5.3. Anson

Mk.1

Porous-Suction

Flight

Test

(1948-1950)

0.1 percent of the free-stream value. The model had a

42-in. chord and had the first pressure jump at 20 percent chord. The first suction slot was put at 5 percent

Based on porous-suctionLFC wind tunnel experiments by Kay (1948), Head (1955) used an AnsonMk. 1 aircraft to test the porous concept in flight tests.The goals of the study were to study laminarboundary-layerflow with uniform suction distributions for zero and adverse pressure gradients, to determine the minimum suction required for laminar flow,and to determine the effectivenessof suction in con-

chord to control possible disturbances caused by simulated debris. The last 19 percent of the model was aflap covered with a sheet of porous bronze mesh forsuction control. Laminar flow was easily achievedback to the flap (81 percent chord). When a flap alignment problemwas corrected,laminarflow wasobserved back to 98 percent chord. These initial lowReynolds number wind tunnel results provided a proofof concept for the slot-suction concept with a pressurejump and verified the idea that at a pressure jump allfluid having a velocity pressure less than the prescribed pressure rise must be removed from the flowfor boundary-layerstability.

trolling transition

induced by roughness

and waviness.

The test section was a 2D symmetric

airfoil

cov-

ered with a porous nylon material (120-mesh phosphor

bronze gauze) covering the suction box. In testing theconcept, the results demonstratedthat laminar flowwas achieved at all rates of suction; turbulent flow wasfound on the same test section with no suction (generated by covering the suction area with an impermeablepaper). For high rates of suction, loss of laminar flowoccurred (in some cases), probably because of surfaceimperfections.Finally,Headshowedthat smallamounts of distributed suction were ineffective in pre-

The success of the wind tunnel experiment

led tothe developmentof a high Reynolds number airfoil.The new airfoil (DESA-2) had laminar flow designedto a chord Reynolds number of 50 x 106 using what ispresently known as the N-factor correlation method(normallyattributedto Smith1956; SmithandGamberoni 1956; and Van Ingen 1956). Note that theearlier document(Smith 1953) was classified until

The design of the 69-slot glove was based on the

pressure and suction distributionmeasuredon the12-slot glove. However, a variation in the hole sizesfor each slot accountedfor the different pressurelosses of the sucked air resulting from a variation inthe chord pressure along a chamber. The slot widthswere selected to balance a local decelerationof theflow due to wide slots (potentiallycausing prematuretransition)and high flow velocitiesin narrow slots(causing unnecessarypressure losses). Furthermore,the issue of surface waviness was controlled by polishing the surface until the wavinesswas reduced1/3000 in/in (height-to-lengthratio) or less.

study. Laminar flow was achieved and maintained in

was made to minimize

the drag by varying the suction distribution. Unlike the

drag rise with maximum chord Reynolds number forthe 12-slot configuration,no drag rise was realized inthe 69-slot test. Groth et al. (1957) postulated that theincrease in drag for the wider spaced slots could becausedby the amplificationof three-dimensionaldisturbances(crossflowand/orGtirtler)or twodimensionaldisturbancesthat may have locally beenamplified between the slots. If the drag increase wasdue to crossflow disturbances,then stronger suctionwould be required at higher Reynolds numbers; thiswould result in increased suction drag and wing profile drag. In addition, the flight tests showed that lowerMach numbers (reducedflight speeds) caused anincrease in lift coefficient, a forward shift of the pressure minimum, and, therefore, a loss of 100 percentlaminar flow. For flights conducted at high subsonicMach numbers (=0.70), regions of local supersonicflow on the glove limited the desired 100 percent laminar flow. For local Mach numbers greater than 1.10, itwas not possible to maintain laminar flow back to thetrailing edge of the test section.

See section 4.1 for a discussion

nation avoidanceflight test.

plane and a glove with 69 suction slots. The justification for the additional slots was that such a multipleslot configurationwould be applicable to an actual airplane wing (i.e., the distance between slots should beminimized to avoid premature transition to turbulencein a high chord Reynolds number flow).

the slot-suction

LFC F-94A

PfenningerandGroth(1961)discussed an 81 slot-suctionexperiment

additionallywhich used

the39

69-slot

during

of insect contami-

approach

with

12 additional

slots

(and

4 chambers)in theregionof

8 to 41 percent chord. For

higher Reynolds numbers, the 81-slot configurationhad a drag increase compared with the 69-slot configuration; however, at lower Reynoldsnumbers andhigher lift coefficients the drag was less than the previous 69-slot test.

5.6. LaterTests

Subsonic

Slot-Suction

Wind

Gross (1964) reported the results of experiments

that were conductedin the NORAIR 7- by 10-FootWind Tunnel using a 17-ft chord, two-dimensional,4-percent-thickslot-suctionlaminar flow airfoil. Onehundred suction slots were located from 1 to 97.2 percent chord. The spanwise extent of the slots reducedfrom 77.4 in. at the first slot to 15.2 in. at the last slot.Full-chord laminar flow was achieved up to a chordReynolds number of 26 x 106. It was suspected thatthe wind tunnel flow quality contaminated the laminarflow for larger Reynolds numbers.

Tunnel

(1958)

Carmichaeland Pfenninger(1959) reported theresults of slot-suction LFC wind tunnel experimentson a 30 swept-wing model. The tests were carried outin the University of Michigan 5-Ft by 7-Ft and theNORAIR 7-Ft by 10-Ft Low-TurbulenceTunnels withthe goal of determining whether surface waviness wasmore critical on swept suction wings compared withunsweptsuctionwings.PreviousresultsbyPfenninger, Gross, and Bacon (1957) and by Bacon,Tucker, and Pfenninger(1959) obtained full-chordlaminar flow to the trailing edge of a swept wing with93 suction slots for LFC. The model had a 7-ft chord

Bacon, Pfenninger,

and Moore (1964) reported

the

experimentalresults of (1) a 4-percent-thickstraightlaminar suction wing and (2) a 30 swept, 12-percentthick, 7-ft chord laminar suction wing in the NORAIR7- by 10-Foot Wind Tunnel to investigate the influence of sound and vibration on the laminar flow extentachieved with LFC suction through slots. Naphthalenesublimation pictures showed that the introduction ofsound for the swept wing resulted in transition in theflat pressure region of the wing and the appearance ofcrossflow vortex signaturesprior to transition. Thestraight wing resultsindicatedthat the frequencydepend,.'nce of transition and sound correlated with thetheory _br Tollmien-Schlichtingwaves. For vibration,additional suction was required to maintain laminarflow.

and the tunnels operated at unit Reynolds number per

foot of 1.7 x 106 or a chord Reynolds number ofapproximately12 x 106. The surface waviness of themodel was 1/3000 in/in, and suction slots were locatedfrom 0.5 to 97 percent chord. Fairings were applied atthe tunnel walls to remove three-dimensionaleffects,and an angle of attack of 0 was imposed on the testarticle. The F-94A flight test parameters were used toguide the wind tunnel experiment. Sine-curve waveswere constructedof Reynolds Wrap aluminum foiland layered using silicone adhesive. The experimentswere conducted with the slots covered by the waves(foil). The results showed that waves of differentlength become critical when h2/_. is a constant (consis-

Gross and Bossei (1964) discussed the experiments znd theoretical analysis of a LFC slot-suctionbody of revolution. The experiments were conductedin the NORAIR 7- by 10-Foot Wind Tunnel, and the30 swc pt-wing model had 120 suction slots. The suction slo:s were connected to 13 suction chambers. The0.003-in. slots were spaced 2 in. apart from 4.84 to75 percent of the model length and were spaced0.5 in. :'rom 75 to 100.4 percent of the model. (Note,100.4 p,'rcent of the model indicates that the last slot

tent with the work of Fage (1943) and the F-94A flighttest results). From the database, the critical wavinessfor swept laminar suction wings was defined as outlined in section 3.2. However, from the limited results

was par:ially positioned on the sting.) Laminar flow to

a lengtt Reynolds number of 20.1 106 was realizedwith the. LFC. The theoretical analysis was comparable wit)l the experiments;however, some disagreement was found because the experimentscould notattain tl:e pure axisymmetric-symmetricflow assumedin the tl"eory.

it appears that multiple waves have smaller allowable

wave ratios than single-waveallowables. Finally, bysealing some of the slots, the slot spacingwasincreased from 0.55 percent (0.4 in.) to 2.2 percent(1.6 in.) chord to determine a measure of sensitivityfor more practical applications.No significant difference in the results was observed in the experimentswith fewer slots.

by using suctionto achievelaminarflow, theReynolds number was 9 x 106. Drag increased as the

Tunnel

Reynolds number was increased. For a Mach number

of 3.0, the test article with no suction had laminar flowfor a Reynolds number of 4.5 x 106. With suction, the

Virtually all the wind tunnel and flight test experiments relating to LFC were conducted in the subsonicflow environment.However, there are a few unclassified supersonic

LFC-related

model was connected to a cylinder to form a total

model length of 40 in.) For a Mach number of 2.5, thedrag without suction was 1.35 times the friction dragof a laminar flat plate and the flow was laminar to aReynolds number of 6 x 106. To recover the same drag

same drag could be achieved

of 6 x 106.

with a Reynolds

number

wind tunnel experiments.

A single-slot,tested at a Mach

Groth (1961) reported the results of supersonic

LFC slot-suction wind tunnel experimentsconducted

9.25-caliberogive cylinderwasnumber of 2.9 in the 8-Inch by

13-Inch Supersonic Blow-Down Tunnel at the University of Michigan to study the flow physics near a slot.Boundary-layerprofiles were measured ahead and aftof the slot with a total-pressuresurvey. A discussionwas given by Groth of the local Mach number andpressure variations near the slot and its impact. Shockwaves emulating from the suction slot increased thesuction drag by approximately10 to 15 percent. Groth(1961) suggested that the installationof many fineslots would reduce this shock-induceddrag.

during 1957 and 1958. Groth, Pate, and Nenni (1965)

reported the results contractedto Northrop Aircraftfrom the U.S. Air Force through 1965. The first studywas conducted in a supersonic wind tunnel at the U.S.Navy OrdinanceAeronauticalLaboratoryin Texas.The model was a biconvex, 5-percent-thick,20-inchord two-dimensionalairfoil. Tests were run forMach numbers of 2.23 and 2.77. Between 23.5 and90 percent chord, 19 slots were cut in the model withsuction extracted into four chambers. The spanwiseextent of the slots decreased from 6.28 in. for the first

Groth (1964a), Jones and Pate (1961), and Groth,

Pate, and Nenni (1965) reported on experimentsconducted in 1961 in the 1-m 1-m (40-in. 40-in.)

slot to 2.56 in. for the last slot, corresponding

to the 8 taperconsistentwith observedturbulentwedgespreading angle. Pressure orifices, thermocouples,andboundary-layerrakes were used for the measurements.Boundary-layermeasurementswere made for severalsuction distributions.For the preliminary tests with nosuction, transition occurred at 40 and 30 percent chordfor Mach numbers of 2.23 and 2.77; this resultedin transitionReynoldsnumbers of 5.1 x 106 and3.9 106, respectively.With the suction model, shockwaves were observed originating from each slot. Thestrength of the waves increased with increased suction. Laminar flow was observed at an angle of attackof 0 for the suction distributions used.

supersonic tunnel at Arnold Engineering and Development Center. A fiat-plate model with a 41-in. chord,40-in. span, and 76 spanwise suction slots was used ina Mach number 2 to 3.5 supersonic flow to study thefeasibility of LFC for supersonic flows. The slot widthranged from 0.004 in. in the front to 0.005 in. in therear of the model. Below the slots, 0.2-in-deep holeswith diameters of 0.042 to 0.062 in. were drilled0.25 in. apart. The instrumentationcould measure surface pressureson the model, suction chamber andmetering box pressures, and temperatures. A rake waspositioned at the rear of the model to determine thestate of the boundary layer. For Mach numbers of 2.5,3.0, and 3.5, full-chord laminar flow was observed toReynolds numbers of 21.8 x 106, 25.7 x 106, and21.4 x 10 6, respectively(up to the tunnel limit). The

Groth (1961) noted that additional tests at Mach

numbers of 2.5, 3.0, and 3.5 were conducted in 1958in tunnel E1 at Arnold Engineering DevelopmentCenter (AEDC) in Tennessee. A 20-caliber ogive cylinder,3.25 in. in diameter (maximum) and 14.443 in. long,was used for the model; 16 suction slots were located

resulting reduction in skin friction drag of 28 and

between 5 and 22 in. of the cylinder with 4 slots connected to one chamber.(Note, the ogive cylinder

These

TS-disturbance41

laminarstabilization

flows

were

where

obtained

compressibility

by

helps considerably;crossflow disturbanceswere

absentfromthistwo-dimensionalflow. The measured

however, with suction, laminar flow was maintained.

At higl-:er Reynolds numbers suction could not main-

boundary-layerthickness and wake drag coefficientswere 40 to 80 percent larger than the theoretical datafor the same suction coefficients.This difference maybe attributable to spanwise contaminationin the experiments or the presence and influence of a detachedshock wave from the blunt leading-edgeplate, whichis not accounted for in the theory.

tain larainar flow. The critical roughness heights of

0.001 t,_ 0.002 in. were determinedfor this high unitReynol, ts number.Pate (1965) and Groth, Pate, and Nenni (1965)reported on wind tunnel results of a LFC 9.2-in. cylindrical body of revolution. Suction was applied through150 slots on the model. Laminar flow was observed atMach number 2.5 to a length Reynolds number of42 x 106 and at Mach number 3.0 to a Reynolds number of 51.5 x 106. The total drag at Mach number 3.0

Shock-waveboundary-layerinteractionstudieswere conductedby Greber (1959) at MassachusettsInstitute of Technology and in 1962 by Groth (1964a)at AEDC to determine if slot-suction could be used to

Additional tests were reported by Groth (1964b) at

Mach numbers of 2.5, 3.0, and 3.5, which were conducted in 1961 in tunnel E1 at Arnold EngineeringDevelopmentCenter. A 20-caliberogive cylinder,3.25 in. in diameter (maximum)and 14.443 in. long,was used for the model, which had the same dimen-

high Mach numbers, had the first slot at 1.6 in. aft ofthe leading edge. No laminar flow was observed withthe first model for Mach number 3.5. The second

sions as the 1958 model. An improved suction system

was used and 29 closely spaced suction slots werelocated between 4.5 in. and 18 in. at spacings of0.5 in.; this led to a more continuous distribution of

model (or modified

suction compared with the 1958 LFC model. A totalpressure head rake was mounted aft of the last slot to

20 x 106 for Mach number 3.5. However, the drag

coeffici _,nt was somewhat higher and was presumed tobe influenced by three-dimensionalityin the tunnel.

measure the state of

the boundary layer. Full laminar

Goldsmith(1964) reported results conductedin1963 in the same AEDC tunnel but with a 72 sweptwing model and at flow conditions of Mach numbers

experimentalboundary-layerthickness measurementswere shown to be 22 percent thicker than theoreticalestimates;however, the theory did not account forpotential shock waves emanating from the slots. Additionally, the effect of surface roughness on the laminarflow extent was measured at Mach number 3.0 andunit Reynoldsnumber per foot of 10 x 106. A0.093-in-diameterdisk with height of 0.0035placed at 2.0 in. on the model. With no suction,tion moved upstream from 14 to 12 in. with theness present for a Reynolds number of 6.3

model) had the first slot at 0.76 in.

down from the leading edge. Full laminar flow was

observed for length Reynolds numbers of 17 x 106 forMach number 2.5, 25 x 106 for Mach number 3.0, and

of 2.0 _nd 2.25, giving a subsonic leading edge to a

supercrJticalleading edge. Contouredwind tunnelwall lin:rs were installed to simulate an infinite (twodimensional flow) swept wing. The model had a 10-in.chord perpendicularto the leading edge and a 33-in.chord it the streamwise direction. Sweeping the wingbeyond the Mach angle zeros the lift wave drag; however, this benefit may be offset by increases in induced

(1960-1965)The July 1966 issue of AIAA AstronauticsandAeronauticswas devoted to discussions on the prospects of Laminar Flow Control and the X-21 LFCflight test. This section summarizesthe content ofthose articles (which primarily focused on work byNorthrop and the Air Force Systems Command), theJune 1967 report of the Northrop Corporation(Kosin1967), papers by Whites, Sudderth,and Wheldon(1966) and Pfenninger and Reed (1966), and AGARDreportsby Pfenninger(1965)and FowellandAntonatos (1965), which summarized the X-21A slotsuction flight experimentand the state of the art inLFC aircraft of that era. NorthropmodifiedtwoWB-66 aircraft to incorporate LFC technology on thewings to demonstrate the feasibility and practicality ofthe design, manufacturing,operation, and maintenanceof LFC aircraft systems. Modificationsof the WB-66aircraft included the removal of the original wings andtheir replacementwith LFC slot-suctionwings, theremoval of the engines and replacementwith aftmounted engines, and the installation of LFC suctioncompressorsin pods mountedunder the wings.Figure 36 shows a modified X-21A aircraft.Nenni and Gluyas (1966) discussed the aerodynamic analysis involved with slot-suction LFC design.In the 1960s, the analysis consisted of defining a wingpressure and velocity distribution,followed by calculations of the viscous boundary-layerflow over thewing, then the suction required to stabilize the boundary layer was determined, and finally the slot spacingand size and the suction system were prescribed. Theprocess was iterative until the desired design wasobtained. By establishing the wing geometry, the wingpressures and velocities can be obtained with transonicwing theory. Notably, the pressure isobars should bestraight and constant along the wing span both toallow the suction slot to see a constant pressure and tominimize the boundary-layercrossflow over a large

near the leading edge of the wing. In the leading-edge

region, the chordwise slots were 0.0035 in. wide andspaced 0.75 in. apart and were used to control the flowon the attachmentline. Strong suction was requirednear or on the attachment line so that the momentumthicknessspanwise

Reynoldsdirection,

number did not exceed 100. In the

the slots were varied in width so

wing of 1 to 5, 5 to 40, and 40 to 100 percent chord,

respectively. The flow passed from the slot in the skinthrough the holes in the structure below the skin, to theduct via the plenum chambers beneath the slots, andthrough the plenum ducts and flowmeternozzlesthrough the inner skin. These slot plenum and holeswere designed to provide a uniform suction distribution along the suction slot to minimize the potential fordisturbances.For the X-21A suction system, 96 suction control valves were employed to independentlycontrol the suction in each slot. The airflow rates forthe system were operational from 85 to 130 percent ofthe designed nominal flow rate to provide variations tovalidate the unproven method for estimating the airflow. For example, the flight condition at an altitude of43 000 ft and a Mach number of 0.75 had airflow ranging from 1.94 to 7.18 lb/sec. For the theoreticaldescription of the suction system involving a continuous distribution,the flight-observedand theoreticallypredicted suction over the wing chord agreed reasonably well except for the lower surface outboard region.Whites, Sudderth, and Wheldon (1966) showed thatfor a Mach number of 0.74 and altitude of 41 400 ft,the fliglat measured and predicted suction distributionagreed in shape but differed in level by 50 percent,with theory underpredictingthe requirements.

To measure the local state of the boundary layer,

total-pressurerakes were mounted at the trailing edgeof the wing. Single probes were positioned at a heightslightly above the laminar boundary-layerthickness.When the state of the boundary layer was laminar, theprobe rt corded a full free-stream total pressure; otherwise, a _maller pressure was recorded due to the probebeing immersedin a turbulent boundary layer. Therelationshipbetween the pressure loss and the transition loc ltion was made both analytically and in flight.Probes were used to measure velocity fluctuationswithin tae boundary layer. Microphones mounted withdiaphra_;ms flush to the surface were used to measureboth velocity fluctuations and to determine sound levels abox e the wing.

Corcerning

the

issue

of allowable

or tolerable

waviness and roughness, the report (Kosin 1967) documents the flight condition of a Mach number of 0.8

that the velocity would gradually be reduced to zero as

the end of the slot was reached to minimize the

and an altitude of 45000 ft, the permissible

stepheightswere0.02 in. for forward-facingsteps,0.009 in. for rearward-facingsteps, and 0.25-in.

potential for vortex formation there. Typical values

of the slot spacing/widthinclude1.1/0.003-0.004,2.0/0.006-0.007,and 1.2/0.005 in/in for regions on the

1965, laminar flow was realized

and 59 jpercentchord for Reynoldsnumbersof20 x 10u, 30 106, and 40 x 106, respectively. TheX-21A programcompletedmore than 200 LFCflights. Figure 37 shows sample results obtained during the flight test for a Mach number of 0.7, altitude of40000 ft, and a chord Reynolds number of 20 x 106;74 percent of the upper surface and 61 percentlower surface had laminar flow.

of the

See section 4.2 for a discussion of the impact of

cloud particulate on laminar flow during the X-21Aflight test.UsinganalysisReynoldsthan 100.wall led

criteria from previous

experiments,therequiredthatthe momentum-thicknessnumberon the attachmentline be lessThe second derivative of the velocity at theto momentum-thicknessReynolds number

correlations for both tangential and crossflow instabilities. Although suction was applied in discrete steps(slots), the calculatedsuction requirementsassumedcontinuous suction on the surface. The suction systemshould be designed to keep slot Reynolds numbersbelow approximately100 to prevent the generation ofdisturbancesby the slot flow. With the suction flowrate determined from boundary-layerstability considerations, the pressure drop through the skin must beset to obtain the desired flow rate.

6. Laminar

longitudinal dynamic motions may require more stringent artificial damping than the minimum acceptablerequirements on the turbulent aircraft. However, bothmotions are of sufficient duration that the pilot corrective action can be applied and the aircraft dynamicsdoes not present a danger to flight safety.

number of 20 x 106. During

laminar flow region was extended to 70 percent chord

at that Reynolds number and from 30- to 55-percentchord laminar flow at a Reynolds number of 30 x 106.

Flow

Control

After

OPEC

Oil EmbargoBecause of the impact of the OPEC oil embargoon fuel prices in the United States in the 1970s, theLaminarFlow Control project (under the NASAACEE Program) was formed to help improve aircraftcruise efficiency. The major NLF and LFC projects inthe United States included various general aviation

An interesting conclusion from Kosin (1967) suggested that future studies should seek to reduce theboundary-layerdisturbanceswhich are generated inthe wing-nose region of the aircraft.

For the flight tests beginning in 1963, the results

showed progressivelyincreasingregions of laminarflow, culminatingat the end of the year with nearly60-percent-chordlaminar flow at a mean aerodynamic45

6.1. Boeing(1977-1978)Kirchner

Research

Wind

(1987)

discussed

Tunnel

LFC

spottim;ss on transition, antiturbulence

screens, honeycombs, and a sonic choke were employed in the 8-ftTPT. Tlae level of ulU,o dropped to between 0.03 and0.06 percent. To simulate an infinite wing flow, upperand lower tunnel wall effects were removed by installing foam wall liners. Figure 38 shows a sketch of theswept-wing model and wall liners installed in the 8-ftTPT wind tunnelwith the anticipatedturbulentregions.

Test

a slot-suction

LFC

swept-wingexperimentthat was conductedin theBoeing Research Wind Tunnel. The principal goals ofthe test were to demonstratethe functionalityof thesuction system, to establish the required suction distribution, and to explore the sensitivity of the flow tosuction level. A 30 swept-wingmodel with a 20-ftchord was designed with slot suction over the first30 percent chord for the upper surface and the first15 percent chord for the lower surface for the designconditionof Mach number 0.8. Confidencein the

Bobbitt et al. (1992) expanded on the discussion to

include the design of the tunnel liner, swept LFC wingmodel, and the type and location of the instrumentation. For a 7.07-ft-chordmodel, the airfoil design hada 12-percent-thick23 swept-wing model, Mach num-

design and analysis tools and the experimental

ber 0.82, C L -- 0.47, and a chord Reynolds number of

20.2 106. In the design of the LFC model, CF disturbances were kept small to prevent CF-TS disturbanceinteractions because the linear design theory could notaccount for nonlinearinteractions.To optimize thedesign, many iteration cycles were required consistingof computing the mean-flow fluid dynamics and the

Tunnel

In 1975, Werner Pfenninger devised a wind tunnel

experiment to determine the impact of a large supersonic zone on a supercriticalwing (concept byWhitcomb and Clark 1965) and application of suction(slotted and perforated) LFC to control the boundarylayer stability characteristics(Bobbitt et al. 1992).

boundary-layerstability properties for specified suction levels. The SALLY (Srokowski and Orszag 1977)and MARIA (Dagenhart1981) boundary-layerstability cod,:s were used for the analysis. For all calculations, distributed suction over 1.5 to 25 percent chordwas enforced with Cq = -0.00015. For the design, anadverse pressure gradient existed to about 25 percentchord followed by a favorable gradient. The modelhad suction capability to 96 percent chord on the uppersurface and to 85 percent chord on the lower surface,with di_'ferent pressure gradients providing the potential for _tudying both TS and CF disturbances.Partialchord saction coupled with the favorable pressure gradient prevented the CF disturbancesfrom growingbeyond N = 4. The TS disturbances grew to N = 10.36at 70 percent chord. A chief concern of the design process wzs the supersonic bubble height limitation (distance b_tween model and tunnel wall) and the desirefor stab le upper surface flow.

The tunnel of choice during 1976 was the Ames

12 Foot Pressure Tunnel because of its good flowquality, demonstratedby the previous achievement offull-chord laminar flow on a swept wing. (See Gross,Bacon, and Tucker, 1964.) However, funding commitments to make flow-qualityimprovementsto theLangley8-Foot TransonicPressureTunnel (TPT)changed the preferred tunnel to the 8-ft TPT in 1978.In the 1980 time frame, the scope of the experimentwas modifiedfrom slot suction only to include aperforated-suctionpanel, and in 1985, the plan wasmodified to include the LFC capability with suction onthe first 20 percent chord of the model. The first testwith a slot-suction model began in 1981 and ended in1985; perforated-suctiontesting began in 1985 andended in 1987; the HLFC test began in the winter of1987 and ended in 1988.

Brcoks and Harris (1987) noted that, for the slotsuction

LFC test, full-chordlaminarflow wasobtaine, t on the upper and lower surface for a Machnumber of 0.82 and a chord Reynolds number of12 106 (unit Reynolds number per foot of approximately 1.7 106). The sonic bubble associated with

Harvey and Pride (1981) discussed the design of

the LFC suction system and required modification tothe tunnel. To minimize the impact of wind tunnelfree-streamturbulencevorticity, noise, and thermal

pressible TS-disturbanceanalysis showed that growthof the disturbances occurred over the first 15 percentchord and suggested that N = 10 would correlate withthe observed transition location. Over the Mach num-

ofat

wing hadReynoldsnumbersthe upper

ber (less than 0.7) and Reynolds

numberrange,N-factors correlated with the experiments ranged from8.5 to 10.5 for TS disturbances.IncompressibleCF-disturbanceanalysis showed that over the samerange the amplificationof the disturbancedid notexceed N = 2.5; this indicated that the transition process on the wing was primarily TS-disturbancedomi-

surface moved upstream to about 80 percent chord and

to about 65 percent chord as the chord Reynolds number approached 20 l0 6. On the lower surface, transition moved to about 75 and 30 percent chord forReynolds numbers of 13 x 106 and 15 x 106. A total

nated. At a Mach number of 0.82 and a Reynolds

number of 20 x 106, TS disturbances achieved N -- 10to 13 at the measuredtransition location of 20 to

HLFC test case,

close to the leading edge, N = 10.5 correlated with the

measurements,and if transitionwas observedatgreater than 40 percent chord, N = 7 correlated withthe measurements.(Section 3.5.3 indicated that higherN-factors are realized for transition in the leading-edgeregion of a wing if the surface curvatureis notincluded in the N-factor calculation.)For a chordReynolds number of 20 x 106, shock interference prevented any meaningful correlation. For the compressible analysis of TS disturbances,N-factors rangedfrom 5 to 7.5 for a Mach number of 0.82, a chord Reynolds number of 20 x 106, and suction applied only upto 10 percent chord. In conclusion, Berry et al. (1987)found transition to be TS-disturbancedominated with

of

HLFC was attempted simply by progressively

turningoff suction over the rear portion of the model until suction was only applied near the leading-edgeregion.For a chord Reynolds number of 10 106, full-chordlaminar flow moved to 53-percent-chordlaminar flowusing suction only in the first 25 percent chord. At achord Reynolds number of 15 x 106, the influence ofchordwise suction extent on the amount of laminarflowafterflowfrom

chord. For this simulated

suction was applied only in the first 8 percent chord.

For CF disturbances,N = 4.5 was reached in the first5 percent chord followed by decay; hence, because theCF modes were decaying at the measured transitionlocation, it was concluded that transition was causedby TS disturbances.For a Mach number 0.82 and achord Reynolds number of 10 x 106, figure 41 shows

location is shown in figure 39. Increasing the Mach

number had a stabilizing influence on the boundarylayer instabilities and the transition location moveddownstream, except at Mach number 0.811 where thetransitionlocation moved upstream.Bobbitt et al.(1996) noted that a significant change in the pressuretook place near Mach number 0.8, which caused dramatic alterations. These alterations may be due to thesupersonic bubble contacting the wind tunnel wall.Using

The compressibleboundary-layerstability codeCOSAL (Malik 1982) and the incompressibleSALLYcode (Srokowski and Orszag 1977) were used to analyze TS disturbancesand MARIA (Dagenhart1981)was used to analyze CF disturbances to correlate computed N-factors with the observed transition locationson the slot-suction wing model. For a Mach number of0.6 and a chord Reynolds number of lO 106, incom-

is shown in figure 40. The results indicated that

about 15 percent chord, the extent of laminarsignificantlyincreased with additionalsuction15 to 20 percent chord.

incompressibleanalyses correlating N-factors of 9 to11 and compressibleanalyses correlating N-factors of5 to 6. They also noted that the N-factor tool should be

flow was achieved

resultstunnelAs describedby Etchberger(1983) and Lange(1984, 1987), the Lockheed LFC concept consisted ofa fiberglass-epoxysubstructure enclosing ducts whichprovided air passage for 27 suction slots. Shown infigure 43, the titanium skin had each slot cut to a widthof 0.004 in. The holes under the slots were 0.03 in. in

were

Full-chord

for theReynolds

Up to 60 percent total drag reductions

achieved for slot-suctiontest comparedunswept turbulent baseline

diamet_;r and centered 0.2 in. apart. Suction was provided by a centrifugal air turbine compressor mountedinside the aircraft. The suction slots covered the uppersurface back to the front spar (12 percent chord). Inthe leading-edgeregion, six slots served both tocontrol the flow and to providefluid for insectcontaminationand ice-accumulationprotection.A60/40 mixture of propylene glycol methyl ether andwater was expelled through the slots. After climb outto 4000 ft, the fluid ejection system was purged fromthe slots. The suction system and glove geometry weredesigned by using computer simulations and wind tunnel experiments.The constructionof the test article

werewith

Suction mass flow required to maintain laminar

flow to 60 percent chord on the upper surfacewas twice as high as predictedfor free-airconditionsSuction over less than 20 percent chord causedtransition to move rapidly forwardThe drag coefficient increased as Mach numberincreased until Mach number 0.82 to 0.825 was

required numerous

reached, when an abrupt increase in laminar

flow was observed (probably due to choking ofthe tunnel and decreased noise).

Flight

trial and error steps.

ture. Fifteen flutes were used to extract air through

0.0025-m. holes spaced 0.03 in. apart. Suction wasapplied from just below the attachment line back to thefront spar. A Krueger shield was used at the leadingedge to deflect or block insects. TKS anti-ice systemwas used on the Krueger shield, and a spray nozzlesystem was appended to the back of the Krueger shieldas a backup system for anti-insect and anti-ice protection of the leading edge pending a Krueger systemfailure, rhe Krueger shield was retracted after reach-

ronmenton boundary-layerreceptivityandtransition for more accurate prediction of suction level requirements for LFC and HLFC

Details of the flight experiment

The Leading-EdgeFlight Test (LEFT) on theNASA Jetstar (Lockheed C-140) aircraft was an element of laminar flow technologywithin the ACEEprogram. The Jetstar flight experiment had objectiveswhich included addressingLFC leading-edgesystemintegration questions and determining the practicalityof the LFC system in operationalenvironmentsviasimulated airline operations.Douglas Aircraft Company and Lockheed-GeorgiaCompany designed andconstructedleading-edgetest sections for the Jetstarright and left wings, respectively. An illustration of theaircraft with suction gloves is shown in figure 42.

ing an altitude of 6000 ft, with the goal of leaving

insect-free leading edge for cruise flight.

an

Both LFC test articles were 61.25 in. long (20 percent of the spanwise extent of the wings) and extendedfrom th_ leading edge to the front spar. At the end ofthe test article at the front spar, both designs had afairing which was used to continue the contours of thetest articles back to 65 percent chord. The contourswere48

altitude of 32000 ft, 97 percent laminar flow was

observed on the Lockheed glove. At the design Machnumber of 0.75, only 74 percent laminar flow wasrealized.See section 4.2 for a discussion

of the influence

of

ice-particulateon laminar flow for the Jetstar flighttest. Note, that the aircraft encounterwith cloudsshown in figure 45 lasted on the order of minutes andthat laminar flow was regained within a few secondsafter exiting the cloud.In addition to demonstratingthat the LFC systemscould be packaged in the leading-edgeregion, laminarflow could be obtained through the suction LFC systems, the simulated airline service demonstratedtherobustness of the LFC systems under normal operatingconditions of typical commercialaircraft (Maddalonand Braslow 1990). As Warwick (1985) noted, theX-21 program had difficulty keeping the LFC systemfree from insects and dirt or dust accumulation.TheJetstar overcame this difficulty by using a Kruegerflap on the right wing and by applying a thin layer offluid on the left wing during takeoff. As a demonstration of the concept, the Jetstar aircraft operated out ofAtlanta,Georgia;Pittsburgh,Pennsylvania;andCleveland, Ohio, and into many other airports in theUnited States in 1985 and 1986 (MaddalonandBraslow 1990). In this service, the aircraft was keptoutside and exposed to the weather (e.g., rain, pollution). Results of the simulated airline service showed

system.

that no operationalproblems were evident with theLFC systems, no special maintenancewas required,and LFC performance was proven through the realization of laminar flow on the test article.

For the Douglas

article,laminarflow wasobserved back to 83 percent of the article length fordesign conditions and back to 97 percent for the offdesign condition of a Mach number of 0.705 and analtitude of 38 000 ft. Powell (1987) and Morris (1987)

primarily been on NLF, mention was made of a LFC

flight test that Cessna and Rohr Industries conductedin August and September1986. The nacelle lengthwas extended by 10 in. and the first 40 percent of the

was not necessary.

49

nacelleon

a Citation III was reskinned

with a woven-

from the fuselage would spill onto the attachment line

and destroy the potentialfor laminar flow. Threedimensionalcalculationswere conducted to theoreti-

wire porous surface called DYNAROHR.

The surfacepressures and boundary-layertransition locations weremeasured. Peterman did not discuss the LFC flight testresults in his presentation.6.5. Dassault

Falcon

50 HLFC

Flight

cally optimize a bump (Gaster 1965) to avoid the turbulent

contaminationproblem.Thisbumpwasdesignedand constructedfor the attachment-lineregion near the fuselage-wing juncture and tested in awind tunnel. Results from the wind tunnel study of asimplified model showed that the bump enabled largerReynolds numbers prior to turbulence onset. A bumpwas manufacturedfor the Falcon 50 aircraft.

Tests

(1987-1990)BulgubureandArnal(1992)and Courty,Bulgubure, and Arnal (1993) noted that the purpose ofthe flight tests on the Falcon 50 aircraft (fig. 46) wasto acquire data to validate and improve design toolsand to show the feasibility of the laminar flow conceptin flight conditions covering a range of Mach number,Reynolds number, and sweep angle to a future laminarbusiness aircraft. The project took place in two flighttest phases plus a wind tunnel validation phase.The first phase (1985-1987)

As shown in figure 46, the installed instrumentation package included (1) 3 rows of static-pressuretaps embedded in the suction article between the flutesto measure the pressure distribution12 hot films each for transitionmounted

of the suction

article,

(3) a series of 14 hot-film sensor arrays on the upper

surface and 14 hot films oriented spanwise on theattachrr_.ent line for attachment-lineboundary-layerstate detection(used only during the leading-edgetransition-contaminationmeasurementsand removed

aimed to demonstrate

that a wing could fly with NLF (optimized airfoil for

extended regions of laminar flow) and to determinethe limits of this concept. The results of the programshowed that transition criteria had been correlated and

for flight tests with suction), (4) a pod installed for

either an infrared camera to record the transition loca-

provided the knowledge required to proceed with the

second phase--aHLFC demonstration.The secondphase (1987-1990)of the flight test aimed to show thefeasibilityof HLFC in a highly three-dimensionalregion near the fuselage. The purpose of the follow-onflight experimentswas to show that laminar flowcould be realizedReynolds numbers

tion or a video camera for recording leading-edge

for a 35 swept wing with flight

ranging from 12 10Uto 20 x 106.The first HLFC flight test phase was conductedinitially without the Gaster bump; the primary objective of the flight investigation was the assessment ofthe TKS anti-icing and insect-avoidancesystem. (Seesection _. 1 for a discussion of the effect of the use of a

The HLFC system was designed

to provideleading-edgeboundary-layersuction aft to 10 percentchord on the upper surface, anti-icing and insect contamination avoidance, and fuselage turbulence contamination avoidance along the attachmentline. Thedesign objective was 30-percent-chordlaminar flow.Shown in figure 46, the perforated stainless steel suction article was placed over the existing inboard wingstructure in close proximity to the fuselage of theFalcon 50 aircraft. The glove was faired into the existing wing with an epoxy resin fairing. Boundary-layersuction was distributed chordwise through six spanwise flutes. In addition, a TKS anti-icing system wasintegrated into the design and performed the additionaltask of insect contamination

TKS an :i-insect system for the flight test.) In addition,

the location of the attachment line was measured forproper placementto the Gaster bump. The secondphase of flight tests was with the bump on the aircraftto deter.nine the effectiveness of the Gaster bump forturbulel ce contaminationavoidance along the attachment line, the effect of sweep angle on the chordwiseextent ef laminar flow, and the effect of suction flowrates an -1distribution on the chordwise extent of laminar flo_. The flight tests were conducted such that thechord Reynolds number variation in the region of thetest article was between 12 x 106 and 20 106. The

leading-edge sweep angle of the test article was nominally 35; however, additional testing was conducted50

at sideslip of 5 which yielded

angle of 30 .

a leading-edge

A 22-ft span segment of the leading-edge box outboard of the engine nacelle pylon and on the left wing

sweep

was replaced with a HLFC leading-edge box as shown

in figure 49. This new leading-edgesection consistedof a perforatedtitaniumouter skin, suction flutesunder the skin, and collection ducts to allow suctioncontrol of the boundary-layerCF- and TS-disturbance

With boundary-layersuction and without thebump, the whole test article was turbulent. For variouscombinationsof Reynolds number and sweep angle,the best case revealed only a very small area of intermittent boundary-layerflow outboard on the test article. With the Gaster bump installed on the leadingedge at 150 mm from the fuselage and with the samesuction rates as in the case of no bump, the boundarylayer was observed to be mostly intermittent. With theGaster bump installed at 300 mm from the wing root,figure 47 shows that most of the test article becamefully laminar. As expected, when the boundary-layersuction turned off, the flow over the test articlebecame

completely

growth from the leading edge to the front spar. The

leading edge included a Krueger shield integrated forhigh lift and insect protection and hot air deicing systems. The wing-box portion of the test area consistedof the original Boeing 757 surface and contour andonly requiredminorclean-up(e.g.,shaved-offexposed rivet heads) to meet surface waviness andsmoothnessrequirements.The design point for theflight tests was Mach number 0.8 at C L = 0.50. Flighttests of many off-design conditions were performed toinvestigate extent of laminar flow as a function ofMach number, unit Reynolds number, and lift coeffi-

turbulent.

The results of this two-phase flight test program

demonstratedthat laminar flow was a viable concept

cient. Flight testing began in February

in August 1991.

for at least the business-type

aircraft. Hence, theELFIN program was established to advance NLF andLFC technologies for subsonic flight. Figure 48 givesa schematic of the range of interest for the projectssupported by the program.

6.6. Boeing

757 HLFC

Flight

As shown in figure 49, flush-mounted

pressuretaps were positioned in the perforated leading edgeand strip-a-tube belts were used to measure the external pressure distributionover the wing box. Hot-filmsensors were used to determine the transition location

Test (1990-1991)

on the wing box and along the attachment

line. Limited infrared camera imaging was obtained and indicated that this technique was useful for boundary-layertransition detection. Finally, wake-survey probes wereused to infer local drag-reductionestimates. The stateof the laminar boundary layer, the internal and external pressure distributions,and the suction system weremonitored in real time onboard the aircraft during the

In the 1980's, it was recognized that conventional

aircraft productionwing surfaces could be built tomeet LFC design constraints. The NASA Jetstar flighttest addressed LFC suction leading-edgesystems anddemonstratedextensive laminar flow in airline-typeoperations. A large, commercial transport demonstration was the natural next logical stage of development.In 1987, NASA, the U.S. Air Force Wright Laboratory, and Boeingated a cooperativetransport aircraft.

1990 and ended

flight test.

CommercialAirplane Group initiflight test program on a Boeing 757

The flight test demonstrated

that the HLFC concept was extremely effective in delaying boundarylayer transition as far back as the rear spar around thedesign point. A sample test condition (fig. 50) showsthat most of the hot films indicated laminar flow

The Boeing 757 high Reynolds number HLFC

flight experiment was designed (1) to develop a database on the effectiveness of the HLFC concept applied

beyond 65 percentchord (Maddalon1991, 1992;Shifrin 1991; Collier 1993). In fact, the suction ratesrequired to achieve laminar flow to 65 percent chordwere about one third of those predicted during the initial design (Maddalon,1991). The wake-rakemeasurements indicated a local drag reduction on the orderof 29 percent with the HLFC system operational,

which resulted in a projected 6-percent drag reduction

for the aircraft (Maddalon1991). However, because

Reneaux and Blanchard (1992) suggested that the

maximum allowableroughnessin the leading-edgeregion would be 0.2 mm and because of this criterion,research should focus on advancing manufacturingtechnologyand insect-impactprevention.Additionally, because convcntionalslats cannot be used in lam-

only about one third of the design suction was required

to achieve laminar flow, significant uncertainty in thedesign tools was a by-product of the flight test. Thisuncertaintyled to the HLFC wind tunnel experimentdiscussed in section 6.13.

6.7. HLFC

ONERA-CERT

T2 Wind

inar flow wings, leading-edge

Krueger flaps or usingsuction to permit higher angles of attack should beexplored for enhancing lift. Finally, the design of theperforated-suctionsystem must focus attention on thehole diameter and spacing, hole pattern and alignment,and the thickness of the surface sheet. The suction

Tunnel

Test (1991)

must be such that premature transition is not induced,

and the pressure drop is such that no outflow isobserved. The hole spacing and size have to be smallcompared with the boundary-layerthickness; a holediameter of 0.06 mm and spacing of 0.6 mm are typical examples of sizes studied.

In 1989, the European Laminar Flow Investigation

(ELFIN) project was initiated and consistedof fourprimary elements that concentratedon the development of laminar flow technologyfor applicationtocommercial transport aircraft. Three of these elementsare related to LFC. These elements were a transonicwind tunnel evaluationof the HLFC concept on alarge-scalemodel, the developmentof a boundarylayer suction device, the developmentof new windtunnel and flight test techniquesfor LFC, and thedevelopmentof improved computationalmethods forlaminar-to-turbulentflow prediction capability (Birch1992).

Reneauxanddesign and testingONERA-CERTT2tion criterionof

rated

To establish criteria for the design of the perfosurface, three tests were carried out in the T2

tunnel. The experiments

studied the critical suctionvelocities for isolated holes, the influenceof holealignment, and validation of the transition predictionmethod. For the experiments,four holes were placedat 20 percent chord and five holes were placed at40 percent chord of an airfoil model with hole diameters which ranged from 0.1 mm to 0.8 mm. Infraredthermographyand liquid crystals were used to detectthe m_nsitionlocation.Criticalvelocitieswere

(1984) was used for the wing design. First, the Airbustransport turbulent wing was modified to achieve thebest compromisebetween transonic performanceandthe HLFC wing. For the wing swept to 27.5 , suction

and

correlated

to

proposed

curve-fit

Square and triangle hole pattern and alignment

were investigated. The critical suction velocities were

was applied from the leading edge to 20 percent chord

and a favorable pressure gradient was maintained to60 percent chord on the upper surface and 55 percentchord on the lower surface. For a Mach number of

larger f_r the triangles; the explanation for the larger

0.82, CL = 0.44, and a maximum chord Reynolds number of 42 x 106, the computedtransitionlocation

Next, hole alignment was investigated

by varyingthe holt alignment to free-stream flow from spanwiseto strea-nwise alignment. With a test section from 17to 34 percent chord, the results indicate that the criticalsuction velocitiesdecreasedwith decreasedhole

ranged from 25 percent chord at the wing root to

55 percent chord at the wingtip for a mean suctionvelocity of 0.1 m/sec. With upper and lower surfacesuction, the computed viscous drag of the HLFC wingwas 45 percent less than the turbulent wing and thetotal drag was 10 percent less than the turbulent wing.Applying suction to the upper surface alone led to aviscous drag reduction of 29 percent and a total dragreduction of 6.3 percent.

spacing The hole spacing seems to have no effect on

transiticn when the distance between holes is 10 diameters. The results also suggested that for hole alignment greater than 30 , the holes behave as though theywere in isolation.

52

6.8.HLFC NacelleDemonstrationFlight Test

(1992)

The flight-test phase of the project extended over

a period of 16 flights totaling 50 flight hr. As shown infigure 51, the HLFC concept was effective over therange of cruise altitude and Mach number and resultedin laminar flow to as much as 43 percent of the nacellelength (the design objective) independentof altitude(Bhutiani et al. 1993, Collier 1993, Fernandezet al.1996). At this transition location, the static-pressuresensors indicated the onset of the pressure recoveryregion, which caused the laminar boundary layer tobecome turbulent. Without suction, significant laminarflow was achieved on the LFC panel; the extent of"natural" laminar flow increased with increasing altitude (perhaps due to passive suction).

The encouragingresults achieved on the Boeing757 HLFC flight experiment and the potential for dragreductionon nacelles led General Electric AircraftEngines (GEAE) to initiate a project with Rohr Industries, Inc., Allied Signal Aerospace,and NASA toexplore the use of LFC on nacelles. The project wasdirected toward the flight demonstrationof the HLFCconcept applied to the external surface of large, turbofan engine nacelles. Bhutiani et al. (1993) stated thatthe main objective of the project was to demonstratethe feasibility of laminar flow nacelles for wide-bodyaircraft powered by modern high-bypassengines andto investigate the influence of aerodynamiccharacteristics and surface effects on the extent of laminar flow.

6.9. NLF

and LFC

Nacelle

Wind

The earlier studies conducted

in the United

States

suggested that significant performance

benefits couldbe realized through the use of NLF and/or LFC onengine nacelles. Before 1991, no flight tests were conducted by the Rolls-RoyceCompany to study LFC;however, wind tunnel tests were conductedwith atwo-dimensionalmodel of a LFC nacelle. The windtunnel test demonstrateda region of substantial laminar flow with sufficient suction. Due to unacceptablelevels of turbulence and noise in the tunnel, the extension of this effort was moved to a low-turbulence9-ft

a turbocompressorunit driven by engine bleed. Forconvenience,the turbocompressorunit was located inthe storage bay of the aircraft. The flow through eachflute was individually metered. The laminar flow contour extended aft over the fan cowl door and was

by 7-ft tunnel at the University

of Manchester.Mullender, Bergin, and Poll (1991) discussed the planto perform a series of wind tunnel experimentstheoretical studies with NLF and LFC nacelles.

accomplishedthroughthe use of a nonperforatedcompositestructure blended back into the originalnacelle contour ahead of the thrust reverser. No provisions were made for ice-accumulationor insectavoidance

Tests

(1991-1993)

A productionGEAE CF6-50C2engine nacelleinstalled on the starboard wing of an Airbus A300/B2commercialtransport testbed aircraft was modified toincorporate two HLFC panels-----one inboard and oneoutboard--asshown in figure 51. The panels werefabricated of a perforated composite material with suction from the highlight aft to the outer barrel-fan cowljuncture. Suction was applied to the surface utilizingcircumferentialflutes and was collected and ducted to

contamination

Tunnel

theoretical studies were aimed at validating

Optimal nacelle designs pointed toward minimizing the length of the cowl to maximize internal performance and drag reductionbenefits. For best highspeed performance, conventionalnacelles have a peakpressure near the lip of the nacelle to distribute thelargest pressure at the most forwardface of thenacelle; the flow was then decelerated over most of the

Static-pressuretaps were mounted on the externalsurface and in the flutes. A boundary-layerrake wasused to measure the state of the boundary layer. Hotfilm gauges were used for boundary-layertransitiondetection. Surface embedded microphoneswere usedto measure noise. A charge patch was used to measurethe atmosphericparticle concentration.An infraredcamera was used for detectingthe boundary-layertransition location. Real-time monitoring and analysisof the state of the boundary layer and suction systemwere accomplishedonboard the aircraft.

nacelle. This pressure distribution produced turbulent

flow over most of the nacelle and a subsequent largeskin friction. Because the circumferentialcurvatureof the nacelle was smaller than the boundary-layerthickness on the nacelle, a two-dimensionalmodelwas used to mimic the nacelleflow. Hot-film,53

total-pressure,and

static-pressuremeasurementsof theboundary layer were made during the wind tunnelexperiment.Using LFC suction, laminar flow wasobserved on the nacelle model. By reducing the levelof suction, TS disturbances were measurable, and withno suction the flow was turbulent. Variations in tunnel

ity (Schmitt, Reneaux, and Pries 1993). The model

had a span of 4.7 m and a mean chord of 1.58 m. Theperforatedleading edge was built into the midspanregion of the wing and had a span of about 0.95 m.Suction was implementedto about 15 percent chordon both the upper and lower surfaces. The titaniumouter skin was 0.9 mm thick and had holes which were

speed indicated that the suction was relatively constant

near the nose over the speed range; however, in themid nacelle region where the pressure gradient wasnearlyflat, notabledifferencesin suctionwereobserved for variation in tunnel speed. The linear calculationssuggestedthat an inviscidinstability(Rayleigh mode) developed and had greatest amplification at 1700 Hz for a tunnel speed of 36 m/secand increasedto 3500 Hz for 60 m/sec.In

correlations achieved 6.6 at a tunnel speed of 36 rn/sec

to 9.1 at 60 m/sec; this indicated that the TS disturbances never evolved sufficient to cause transition.Ratheratransition.

6.10.

VFW

separation

614 HLFC

bubble

developed

Transonic

Wind

As suction was increased

the transition

front

moved aft. Laminar flow was achieved to 50 percent

chord on the upper surface and to 30 percent chordon the lower surface. Data gathered from the test

causing

Tunnel

were usedforand calibrationmethodology.

suctionof the

6.11.

NLF

European

systemlaminar

and HLFC

designcriteriaflowprediction

Nacelle

Test (1992)

Demonstrator

In 1986, the German laminar flow technology program, supported by the German Ministry of Researchand Technology(BMFF),began wind tunnel andflight experiments for NLF and LFC (Redeker et al.1990). Ktimer (1990) noted that part of the programinvolveddetermining(or discriminating)betweenwhen NLF is preferred and when HLFC or LFC is a

In 1992 and 1993, a cooperative program was conducted by DLR, Rolls Royce, and MTU with the goalof investigatingin flight the prospects of achievingextensi,'elaminarflow on aircraft engine nacelles(Barry et al. 1994). The test vehicle chosen for theproject was the VFW 614 ATTAS aircraft which hastwin Rolls-SnecmaM45H turbofans.The placement

more appropriate choice for a particular aircraft. Two

of the major milestones of this program involved NLFwind tunnel tests and flight research on a VFW 614and Fokker 100 research aircraft to gain a database ofTS-disturbanceand CF-disturbance-dominatedtransition for code calibration.

of the nacelle on the aircraft is shown in figure 53. The

program had the usual goals of demonstratingdragreductic_n with NLF and HLFC on a nacelle, verifyingthe de,'ign methodology,verifyingmanufacturingtechniqlles, and validating the anti-insect transpirationsystem.

The successful

VFW

614 and Fokker

100 NLF

Flight

Tests

(1992-1993)

For the NLF portion of the test program, two new

composite nacelles were constructedby Hurel-Duboisfor the program. One nacelle consistedof baselinelines an:l the second nacelle consisted of a new set of

flight tests led to a transonic wind tunnel evaluation of

the HLFC concept, evaluationof wind tunnel testtechniques, and developmentof viable boundary-layersuction devices. In March and April of 1992, a1:2 scale model of one of a VFW 614 wing was builtwith leading-edgesuction and tested in the ONERAS 1MA transonic tunnel--thefirst LFC test in the facil-

aerodynamiclines, conducive to laminar flow. A thirdnacelle was designed for validation of the HLFC concept, which included a liquid transpirationinsect contamination avoidance system. (See Humphreys 1992.)54

Instrumentationto measure the pressure, temperature,and transition location is illustrated on the test section

The second phase of the program involved the

testing of the A320 vertical fin with leading-edgesuction in the ONERA S1MA facility. The 1/2-scalemodel in the tunnel is shown in figure 57. The objectives of the wind tunnel experiment were to simulateflight Reynolds numbers on the model, calibrate thetransition prediction tools, and establish LFC suctiondesign criteria. Finally, Anon. (1995b) reported thatthe A320 HLFC fin flight test program was scheduledto be completed by 1996. (Prior to the publication of

in figure 53. The flight test portion of the program

consisted of about 93 hr which clearly demonstratedthat laminar boundary-layerflow was achievable over60 percent of the nacelle length in the installed environment over a large range of flight conditions forboth laminar flow concepts tested. For the NLF concept, figure 54 shows the design and measured pressures at two radial locations. Very good agreementbetween the computed and observed pressures is realized at t_ = 30; however, significant disagreement wasfound at _ = 140 near the pylon. This disagreementcan be attributed to the computations not including thepylon in the design. Noise and vibration had little orno effect on the ability to achieve laminar flow forthis design.The liquid transpiration-styledinsectcontaminationavoidancesystem was operated successfully during the course of the flight testing.

6.12.Flight

A320

Laminar

Test Program

Fin Wind

Tunnel

the present report, no flight test data were available.)

The development of the A3XX program at Airbus hasallowed for the success of the A320 LFC fin programby requiring the power plants of the A3XX to be positioned closer to the wing and for suction LFC nacelles(Birch 1996).

6.13. LangleyTunnel

and

8-Foot

HLFC

Wind

TransonicTunnel

Pressure

Test (1993-1995)

Although the Boeing 757 HLFC flight test experiment demonstrated

significant runs of laminar flowusing leading-edgesuction, sufficient uncertaintyinthe design tools made the technology an unacceptablerisk for the commercialmarket. To provide a betterunderstandingof the complex physics of flow over aswept-wing geometry, to provide a calibrationdatabase for the LFC design tools, and to better understandthe issues of suction-systemdesign, a joint NASA/

(1993-1998)

Figure 55 shows an illustration of a 1987 plan by

Airbus Industries in close collaborationwith ONERAand DLR to enable LFC capability for subsonic transport aircraft. The program consistedof theoreticalanalysis, a large wind tunnel evaluation, and a flighttest program of the vertical fin of the A320 aircraft(ultimately geared toward the application of laminarflow to wing and tail surfaces of a future advanced aircraft). The vertical fin of the A320 aircraft was chosenas the candidateto test the feasibilityof HLFCbecause of the availability of an aircraft for flight testing, simple installation, no de-icing system, attainmentof flight Reynolds number in an existing wind tunnel(ONERAS1MA at Modane),and minimizedcost(Robert1992a; Redeker,Quast, and Thibert1992;Thibert, Reneaux, and Schmitt 1990).

Boeing HLFC wind tunnel experiment was conducted

in the Langley8-Foot TransonicPressureTunnel(Phillips 1996).A swept-wingmodel with a 7-ft span and 10-ftchord was installed in the tunnel in January 1995 andtests were conducted throughout the year. Tunnel liners were installed to simulate an infinite swept wing.Over 3000 infrared images and 6000 velocity profiles(hot-wire data) were obtained during the test, and thedata were made available to the team of researchers inreal time via encryptedtions (Phillips 1996).

Shown in figure56, boundary-layerstabilityresults indicated that laminar flow is expectedtoapproximately40 percent chord for the baseline A320fin and to about 50 percent chord for the HLFC A320fin (using a reasonable amount of suction). A benefitstudy with the projected amount of laminar flow indicates that an aircraft drag reduction of 1.0 to 1.5 percent is possible by laminarizing the vertical fin.

As statedLFC designdesigns. Thesuction level

World Wide Web communica-

by Johnson (1996), an assessment of the

criteria was made to help guide futureinfluence of hole size and spacing andand distribution on the transition location

was recorded and correlated with the design tools.

Laminar flow was easily obtained back to the pressure

55

minimumface

with sufficient

roughness

underway

suction

to characterize

Detailedfor inclusion

6.14.

and

suction levels. Detailed

level

measurements

the leading-edge

results are not available

sur-

flow to higher chord Reynolds numbers. As discussed

by Gottschalk(1996), such a concept proposed byNorthropGrummanCorporationwouldhaveasharp supersonicleading edge and result in a thinattachment-lineboundarylayer and a very smallmomentum-thicknessReynolds number. Such a flowshould be stable and have a laminar attachmentline.Crossflow disturbances could be avoided with the low

are

panels.

in the literature

in this publication.

High-Speed

Civil Transport

(1986)

wing sweep and, with appropriate wing shaping, a partially NLF wing could be achieved. LFC would berequired on the rooftop of the wing to extend theregion of laminar flow to higher Reynolds numbers.Concerningthe use of thermal LFC, Dunn and Lin(1953) have shown in the early 1950s that cooling canbe used to suppress disturbances. As shown by Boeing(Parikh and Nagel 1990), cooling has a large impacton TS disturbancesand only a subtle influenceonCF disturbances;hence, cooling would not be usefulin the leading-edgeregion of swept wings for CFstabilization.

In 1986, NASA and the U.S. airframe and engine

manufacturersdetermined that the long-range travelmarket was conducive to a supersonic airliner (highspeed civil transport,HSCT); however,significanttechnologicaladvances were required. The advanceswould require an aircraft to fly slightly faster than thespeed of the Concord but with nearly twice the rangeand three times the number of passengers at an affordable ticket price while not damaging the environment.As shown by Kirchner (1987), laminar flow couldlead to significant benefits for a supersonic transport.When consideringthe application of NLF and LFCtechnologiesto the supersonic flow regime, the highcost and limited availabilityof flight test aircraftinhibits the advancementof these technologies.Military jet fighter aircraft, the Concord, and the TupelovTu-144 currently fly at supersonic Mach numbers andare potentiallyviable candidatesto serve the LFCresearch community;however, the design and manufacturing of most of these aircraft were devoid of thefuture potential use for LFC missions and potentiallyhave unacceptablesurface waviness, roughness, andaircraft-specificobstacles. Wagner et al. (1990) presented the status of supersonic LFC through the 1980s.

In contrast to the low-sweep

supersoniclaminarflow concept proposed by Northrop Grumman,thehighly swept wing would have a subsonic leadingedge, a blunt nose, and higher momentum-thicknessReynolds number. As Wagner et al. (1990) noted, theturbulentbaselineHSCTconfigurationsby TheBoeing Company and McDonnellDouglas Corporation were making use of the second approach. Withthish_gh-sweepwing,theissueof turbulentattachment-linecontaminationmust be addressed andsuctionLFC wouldbe requiredto controltheCF-dominatedtransition process in the leading-edgeregion of the wing. For long chords typical of theHSCT configurations,an additional strip of suction (orthermal) LFC would be required on the wing to delaythe TS- dominated transition process.

In spite of these limitations, technology can be

advanced by making use of these aircraft when theyare made available. Toward the goal of advancingNLF supersonictechnology, flight experimentswerecommencedin the United States toward gaining abetter understandingof the viscous flow physics. Asummary of the NLF results for supersonic aircraft arepresented in appendix B.

Williams (1995) noted that a proposed HSCT carrying 305 passengers and flying 5000 n.mi. with 1990technologywould weigh almost 1.25 million lb attakeoff and would not meet the current noise requirements. A technologydevelopmentprogramwouldneed tc reduce the weight by almost 50 percent tomake tae HSCT feasible.Toward overcomingthetechnic;d obstacles, NASA commencedPhase I of a

Two fundamentalapproaches were posed for thesupersonic laminar flow wing. The first approach wasa low-sweep wing which involved the design of a NLFleading-edgeregion and low-suction(or thermal)LFC on a section on the wing to extend the laminar

High-SpeedResearch (HSR) Program in partnershipwith U.S. industry. Phase I focused on developing reliable methods to predict engine-emissioneffects on theozone,

56

noise reduction

technologies,

and the potential

advantages(SLFC).

of

supersonic

laminar

flow

blowdown facility supplied with dry high-pressure

airwhich exhausts into large vacuum spheres to providerun times on the order of 30 min. The nozzle throat is

control

highly polished to maximize the extent of laminar

flow on the nozzle walls. Upstream of the sonic throat,suction was used to remove the turbulent boundary

Feasibility studies by Boeing Aircraft Company

(Parikh and Nagel 1990) and McDonnellDouglasAircraft Company (Powell, Agrawal, and Lacey 1989)were funded to determine the benefits of supersoniclaminar flow control applied to the HSCT configuration. Reductions in gross takeoff weight, mission fuelburn, structural temperatures,emissions,and sonic

layer that exists on the wall. The fresh laminar boundary layer evolved through the contoured nozzle untilthe boundary layer undergoes transition to turbulence.The location of this transition point governs the lengthof the low-disturbancetest-sectionrhombusand is

boom were predicted by incorporating

SLFC technology on a HSCT configuration(see section 2).Because

of the favorable

results

achieved

directly influenced by the unit Reynolds number of the

flow. As the unit Reynolds number increases, the sizeof the quiet test-section rhombus decreases; however,the Reynolds number based on the length of the quiettest core increases. The tunnel was capable of operat-

with

Phase I of the program, HSR Phase II was initiated to

performadditionalresearch toward advancingthestate of technology to make the HSCT economically

ing in conventionaldisturbance) mode.

viable. As part of Phase II, the low-disturbance

windtunnels at Langley and Ames Research Centers andthe F-16XL aircraft at Dryden Flight Research Centerwere used to advance the state of the art in supersoniclaminar flow control. An overview of the understanding of SLFC up to 1987 was providedMalik (1987).

6.15.

Supersonic

LFC

Quiet-Tunnel

noisy

mode

or in quiet

In the SLDT, measured transition

bers were shown to be comparable

(low-

Reynolds numwith transition

observed in flight. Creel, Malik, and Beckwith (1987)

and Creel, Beckwith, and Chen (1987) used the quiet

by Bushnell and

tunnel to study boundary-layer

instabilities on a leading edge of a swept cylinder. The results suggestedthat transition was affected by wind tunnel noise onlywhen large roughness was present on the model, thelocal roughness Reynolds number correlated with thetransition location for a wide range of Mach numbers,and linear stability theory showed good agreement forthe experimentalcrossflow vortex wavelengthof thedominant mode. Morrisetteand Creel (1987) studied

Tests

(1987-1996)Conventionalsupersonicand hypersonicwindtunnels are dominated by acoustic disturbances radiated from the turbulent boundary layers on the tunnelwalls. The emanationof these disturbancesfollow

the effect of surface roughness and waviness on transition in the SLDT. Controlled roughness and wavinesswere imposed in the supersonicflow and comparedwith subsonic correlations.Eight 15-in. long and 5

Mach lines. To study laminar flows (i.e., transition,

boundary-layerinstability, and LFC), the test sectionin the tunnel must be clean (defined as free-streampressure fluctuationsbelow 0.1 percent). This sectionfocuses on the research primarily supported by theHSR project and conducted in the Langley SupersonicLow-DisturbanceTunnel(SLDT)and the AmesLaminar-FlowSupersonicWind Tunnel (LFSWT).For more details about quiet tunnels, refer to the

half-angle wavy cones were tested, where the wavelength of the cones correspond to the most amplifiedTS disturbancefor the smooth cone. A fixed surface

pitot tube was used to measure transition as a function

of total tunnel pressure. Results with wall wavinessindicated that the tunnel running with a noisy environment led to lower transition Reynolds numbers compared with the results in the quiet environment.Also,the results suggested that the transition location was afunction of aspect ratio (wave height over wavelength).The quiet tunnelresultsfor roughnessmatchedwith the correlation by Van Driest and

number(SLDT)

McCauley (1960) for three-dimensional

cones. Morrisetteand Creel (1987)

review of quiet tunnel technology

(1992).

by Wilkinson

et al.

3.5 SupersonicLow-DisturbanceTunnelat Langley Research Center. The tunnel is a57

roughness onconcludedthat

waviness had less effect on transition than a single trip

of comparable height, and the effect of noise on critical and effectiveroughnessReynoldsnumbersappeared small.

raphy was used to assess the state of the boundary

layer oTathe cylinder for variations in free-stream conditions. Observationsindicate that the boundary layerremainedlaminar up to and including the largestattachment-lineReynolds number of 760. Using tripwires to control the state of the boundary layer, theresults suggestedthat the free-streamdisturbanceenvironmentimpacted the transition location; this con-

In support of the F-16XL SLFC flight experiment,

models were developed for the Langley quiet tunnel tocalibrate the design tools for NLF and LFC and tostudy attachment-linetransition. Iyer and Spall (1991)and Iyer, Spall, and Dagenhart (1992) performedlinear stability theory calculations using CFL3D for themean flow and COSAL for boundary-layerstabilityfor the F-16XL leading-edgesectionmodel. The

firmed that designs based on conventional

nels were too conservative.

6.16. F-16XL(1989-1996)

15-in. model had a leading-edge

sweep of 77.1 with anormal Mach number of 0.78. Traveling CF disturbances were found to have the largest amplification;however, distributedsuction was shown to stabilizethe flow so that N = 10 was not exceeded over the

LFC

Flight

Tests

Supersonic LFC flight tests were conducted by a

NASA and U.S. industry team to demonstrate the feasibility of laminar flow in supersonicflight. TwoF-16XL aircraft (XL has delta wings) are on loan toNASA from the U.S. Air Force to serve as testbeds.

The F-16XL wings have inboard sweep of 70 and

outboard sweep of 50 , similar to the proposed HSCTwing configuration.NASA and RockwellInterna-

in figure 58, the calculated N-factors correlated well

for N = 14 over a range of free-stream unit Reynoldsnumbers and angle of attacks for the solid model. Theresults suggested that traveling crossflow disturbancesprobably dominated the transition process. A SLFCporous-suctionmodel was developed and tested butthe results are not available for this publication.

tional Corporation carried out the flight tests with the

In ]990, flight testing began using a suction glove

on the F-16XL Ship 1 (shown in fig. 59(a)). ARockwell-designedperforated-suctionglove was fabricated and installed on an existing wing of Ship 1 assketched in figure 59(b). Because of the geometricalconstraints of implementinga glove on Ship 1 (gloveheight of less than 2 in. above the existing wing surface and 10 in. in front of the leading edge), activesuction was limited to the first 25 percent chord andattachment-lineinstabilities were the primary focus ofthe LFC experiment.Woan, Gingrich, and George(1991), Anderson and Bohn-Meyer (1992), and Norris

At the Ames Research Center, a Mach 1.6 quiet

tunnel was constructedto minimize the free-streamdisturbances.This was accomplishedby using a lowdisturbance settling chamber to produce steady supersonic diffuser flow and low structural vibration andincluded smooth (polished) walls to produce laminarboundary layers on the nozzle and test section. Wolf,Laub, and King (1994) presented results for flow quality and tunnel transitionaspects of this continuousoperation facility. Supporting the F-16XL SLFC flightexperiment, a section of the passive glove was used tostudy the leading edge of the wing. A comparison ofthe surface pressure distributions measured in the tunnel compared well with CFD predictions at an angle ofattack of 0; however, the agreement was rather poorfor flight test measurements.More recent attachmentline transition experimentson a swept cylinder werereported by Coleman et al. (1996) and Coleman, Poll,and Lin (1997) in the Ames tunnel. Schlieren

Supersonic

noisy tun-

(1994) noted that the perforated-suction

glove onShip 1 was designed for a Mach number of 1.6, altitude of 44 000 ft, angle of attack of 2, momentumthickness Reynolds number on the attachment line ofless than 114, and a unit Reynolds number per foot of2.53 x 10 6. No laminar flow was achieved at thedesign point; however, laminar flow was observedoff-design conditions. Figure 60 shows the amount

atof

laminar flow with and without suction for a given

flight test condition; hot-film data indicated laminar

photog58

flow to the outboard portion

and Bohn-Meyer1992).

of the glove

(Anderson

flight tests was to obtain surface pressure data to

calibrate the Euler design codes, particularlyin theleading-edgeattachment-lineregion. Preventingthefuselage turbulent boundary layer from contaminatingthe attachment-lineregion of the wing was a secondmajor technical issue which was addressed in the firstphase of flight tests. The third technical area of interest involved characterizingthe acoustic disturbancefield and disturbanceswhich could come from the

Woan, Gingrich, and George (1991) reported on

the design, analysis,and validationof a coupledNavier-Stokesand compressiblelinear stability theoryapproach for supersonic LFC design. Validation wasobtained by using the methodology to design the suction LFC glove for the F-16XL Ship 1 and then bymaking a comparison with flight-measuredresults. Atechnology goal of the methodologywas to obtain adesign which minimizessuction requirementsandsimultaneouslydefines a pressure which is conduciveto stabilizing the boundary layer. Overall,results were in reasonably good agreement

The perforated-suctionglove for Ship 2 wasdesigned in a collaborativeeffort between Boeing,McDonnell Douglas, Rockwell, and NASA. A photograph of Ship 2 and a sketch of the LFC test article areshown in figure 61. Because of the asymmetryofShip 2 with the suction glove, stability and control ofthe Ship 2 configurationwas tested for safety assurance in a wind tunnel. For the flight article, theperforated-suctionSLFC glove was constructedofinner and outer titanium skin and aluminum stringers.Suction was obtained by using a modified Boeing 707turbocompressor.Norris (1994) noted that suction was

Ship 1 database. Mean-flow results from the NavierStokes codes were used with the COSAL boundarylayer stability code for correlations with the availabletransition Ship 1 data. Stability calculations(for anN-factor of 10) indicated that transition would occur at1.5 in. from the leading edge without suction; shownin figure 60, laminar flow was restricted to very nearthe leading edge in the flight test with no suction. Thecomputationsshowedthree distinct shocks whichmust be tracked for laminar flow management.Theseshocks emanated from the nose, the canopy, and theengine inlet (underneath the aircraft).

appliedthroughsome10 millionholesand20 individual suction regions on the glove surface.Wagner et al. (1990) and Fischer and Vemuru (1991)noted that the F-16XL Ship 2 SLFC flight experimenthad objectives of achieving laminar flow over 50 to60 percent chord on a highly swept wing, of deliveringvalidated CFD codes and design methodology,and ofestablishinginitial suction system design criteria forLFC at supersonicspeeds. The suction glove wasinstalled on Ship 2 and the first flight was conductedOctober 13, 1995. The first supersonicflight tookplace on November22, 1995. The first suction-on

Flores et al. (1991 ) used thin-layer Reynolds averaged Navier-Stokes

equations to study the sensitivityof the attachment line and crossflow velocity profilesto changes in angle of attack for Ship 1. The resultsshowedthat as angle of attack increased(1) theboundary-layerthickness and streamwise velocity profiles had no significantchanges, (2) the attachmentline moved from the upper surface to the lower surface, and (3) the crossflow velocity componentat afixed location on the upper surface of the wingdecreased. This information is important for determin-

supersonic1996.

ing the optimal amount of suction required for a given

position on the wing to obtain laminar flow.In the 1991-1992ments were obtained

time-frame,for the flow

Flight

Research

Center.

January

24,

Similar to Ship 1, Ship 2 had aircraft-specific

shock and expansion waves which influenced the flowon the wings. Althoughcanopy and engine inletshocks spreading out over the wings and expansionwaves from beneaththe wing causeda highlythree-dimensionalflow field and difficulties in obtain-

flight measureon the F-16XL

flight test was accomplished

ing laminar flow on the attachment-line

region at thesame test conditions,significantprogresstoward

The goal of the first

59

accomplishingthe goals was achieved. In spite ofthese test aircraft-dependentobstacles, Smith (1996)noted that the supersonic laminar flow control flightexperiment achieved about 70 to 80 percent of the initial goals.

7. Concluding

Reynohls number, wind tunnel test techniques

forHLFC ,-onfiguration development;the demonstrationof acceptable reliability,maintainability,and operational characteristicsfor a HLFC configuration;andthe ability to predict and guarantee benefits to the airline customers. In 1991, a Senior Vice President of anairframe systems manufacturer stated that before laminar flow control could be used on commercial aircraft,

Remarks

This publication has reviewed

foundationalstudiesand more

the long-term technical and economic viability of the

technology must be demonstrated.Although many ofthese issues have been addressed subsequent to thisstatemeat, the future of subsonic and transonic LFC

some of the early

recent U.S. and

European projects which had goals of solving technical obstacles associated with the application of laminar flow control to advanced transport aircraft. Thetechnologyhas the potential to offer breakthroughimprovementsin aircraft efficiency by leading to significantreductionsin aircraftfuel consumption,extendingrange or increased payload, reductions inemissionsand noise, and increasing cruise lift anddrag, and reducing maximum gross takeoff weight.Much progress has been accomplished toward the goalof commercialincorporationof laminar flow control(LFC) (and natural laminar flow (NLF)) on wings,tails, and engine nacelles.However,becausetheapplication of the technology leads to additional sys-

technology must reside in a large-scale demonstrator

to stud) the long-term reliability of the performanceand flight-safetyoperations,in refined design tooldevelopment,and in the longer term understandingofthe effects of wind tunnel flow quality on the laminarflow (LF) extent. An alternative future resides in thedemonstrationof innovativeLFC control systems.Perhaps, advancesin micro-machine,synthetic-jet,smart-rraterialtechnologieswill lead to orders ofmagnitude improvementsin efficiency, reliability, andcost-effectivenessof these future LFC systems, andLFC will be an integral part of this revolutionarynewaircraft.

tems and some uncertainty in the maintenance requirements and long-term structural integrity due to thesystem, questions still remain which must be resolvedrelative to long-term operational and reliability characteristics of current hybrid laminar flow control(HLFC) concepts before the aircraft industry can guarantee the sustained performance of the LFC vehicle totheir airline customers.

In the supersonic vehicle class, the 1990s brought

the first flight demonstrationof LF achieved by supersonic la_ninar flow control (SLFC) through the successof a NASA-industryteam. In 1990, a General Manager of a major airline company stated in a talk on thehigh-speedmarket in the next three decades that,although the subsonic fleet will play the role of serving the low-yield mass traffic markets, the supersonictransport will be a big part of the intercontinentalfleetof the future. Looking at historical data, the long-rangeaircraft _ntering the market and replacing an existingaircraft aas never been smaller than the aircraft beingreplaced. Based on these data, the smallest intercontinental supersonic transport (SST) will have a capacityof no less than 300 seats (at moderatelyhigher-20 percc nt-----cost than the subsonic cost). The benefitsof LFC increase with the size of the aircraft. If this

The 1980s and 1990s brought the successful demonstration

of a LFC aircraft (Jetstar and Falcon 50LFC flight tests) in airline operationsprevention systems, the achievement

and with insectof laminar flow

at high Reynolds numbers (Boeing 757 HLFC flight

test), the achievementof laminar flow on a HLFCengine nacelle (A300/GE and VFW 614 nacelle flighttests), and various LFC wind tunnel tests (Langley8-Foot Transonic Pressure Tunnel and ONERA S 1MALFC tests). However, from the airframe company perspective, some technology issues exist which requireattention prior to the acceptance of LFC. These issuesinclude the resolution of potential performancepenalties versus projectedHLFC benefits(leading-edgeKrueger versus conventionalleading-edgeslat system); the developmentof HLFC compatibleiceprotectionsystems; the developmentof viable high

subsonic: trend of larger aircraft entering the market

continue s, the LFC technology could be an even moresignificant competitiveadvantage to a next generationairplane.Environmentalissues, materials,systems,engines, and supersonic laminar flow control are someof the research which ought to be pursued for thedevelopment of a supersonic transport.60

The

reduced

unfeasibilitypromise

of the

of benefits

fuel prices.

priority

of LFC

technologybeing

As the cost

residesbut

intimately

increases,

the benefits

nate fuels

are introduced

rather

of LFC

with

any

of reduced

noise

the

supersonicwith LFC.

aircraft)

tied to the aircraft

of fuel decreases

value, the benefits

and hence futuredecreaseto obscurity;conversely,

not with

in real

into the equation,

Even

emissionsremain

dollar

prospectsof LFCas fuelprice

increase.

and

NASA Langley Research Center

Hampton, VA 23681-2199June 18, 1998

if alter-

the benefits

61

(and

attractive

heat

stress

achievements

on

AppendixSubsonicResearch

ANatural

Laminar

control behavior relative to FAR Part 23 and (2) climb

performancedecreased l 0 percent, which was consistent with the increased drag associated with a trippedboundary-layerflow.

cation of contours or waviness, the flight test results

indicated that laminar flow on the wings and empennage was responsiblefor the previouslymeasured

and flight tests. For example,

a CessnaT210Rresearch aircraft was used in the late 1980s to validate

lower-than-expectedzero-lift drag coefficient. No premature transition was observed due to waviness, contour discrepancies,or surfacedents.Significantregions of laminar flow were realized in the slipstreamregion. Insect-debris contaminationin flight indicatedthat 25 percent of the insects caused transition. Thefact that transition was realized downstreamof the

the use of NLF for aerodynamic

performancegains.This research airplane had a NLF wing and horizontalstabilizer and a smoothed vertical stabilizer. The airfoil was designed to achieve 70 percent NLF on bothupper and lower surfaces; this resulted in low drag at acruise Reynolds number of 10 x 106. Murri and Jordon(1987) and Befus et al. (1987) performedfull-scalewind tunnel and flight tests of this aircraft. Under ajoint research program, NASA, Cessna, and the Federal AviationAdministration(FAA) addressedtheflight testing of a NLF aircraft to simulate FAR Part 3certification.Relatedto certification,Manueland

minimum pressure suggests that acoustic, surface, or

turbulence disturbances are not responsible for transition; rather, the amplificationof TS disturbancesorlaminar separationin an adverse pressure gradientdominales the transition process. NLF was achievedon appraximately40 percent of the wing and 50 to70 percent of the propeller. In a comparisonof thewaviness of the BellancaSkyrocketII productionquality with the filled and sanded wing test section ofthe King Cobra (see Smith and Higton 1945), it isclear that the production quality of more modern surfaces has less variation, sufficient for NLF and LFC

Doty (1990) describe the impact of the loss of laminar

flow on the Cessna T210R and make quantitativecomparisons of the ability of the aircraft to meet certification under these conditions. Three test conditionswere explored:

The

Natural

transition

II

Holmes et al. (1983) reported on a flight investigation of NLF on a high-performance,

single-propeller,composite aircraft. The primary goals of the flight testwere (1 } to address the achievability of NLF on a modem composite production-qualitysurface and (2) toaddresssome of the NLF-relatedmaintainabilityissues (e.g., insect contamination).The flight envelopeenables unit Reynolds numbers up to 1.9 x 106 andchord Reynolds numbers of 12 106. Without modifi-

the airfoil had a favorable pressure gradient to about

70 percent chord on the upper surface (crossflow disturbanceswere not consideredin the design). Theglove was installed on the wing to achieve the desiredpressure distributionat l0 wing sweep. The flightresults showed that laminar flow was obtainedto

III

Wentz, Ahmed, and Nyenhuis (1984, 1985) discussed the results of a Langley ResearchCenter,Wichita State University,Cessna Aircraft Co., andBoeing CommercialAirplane Company joint researchprogram on NLF. The study used a business jet aircraft with the following objectives:1.

To determine

the transition

location

56 percent chord on the upper surface at 9 sweep, to

21 percent chord at 25 sweep, with chord Reynoldsnumbers from 23 x l0 6 to 28 l0 6, respectively.Themaximum run of laminar flow on the lower surfacewas 51 percent wing chord at 16 wing sweep to 6 percent chord at 25 sweep (sideslip). The overall resultsfrom the F- 111 TACT NLF flight experiment showedlaminar flow but not as much as expected. Besides not

at various

Mach numbers

and Reynolds

2.

To determinetransition

the effects

numbers

3.

To determinetransition

impact

4.

To check the validity of boundary-layer

of wing

of engine

sweep

on

acoustics

on

attached to both wings. The primary goal of the study

was to demonstratelaminar flow at higher Reynoldsnumbers for swept wings. The glove geometry consisted of a supercriticalNLF airfoil designedbyBoeing and NASA to investigateNLF at transonicspeeds. For the design lift coefficient of 0.5 at a Machnumber of 0.77 and a Reynolds number of 25 x 106,

accounting for potential crossflow-induced

transition,the F- 111 had a limited spanwise extent of test sectionand had a crude method for determining the transitionlocation.

Sublimatingchemicals and hot-film anemometryareused to detect transition. The test section on the wingwas covered with fiberglass and filled and smoothedto minimizeroughness-relatedeffects causedby

designed (Viken 1983) to achieve 70-percent-chord

laminar flow on both upper and lower surfaces at thedesign Reynolds number of 10 106 and Mach number of less than 0.40. In the wind tunnel experiment,laminar flow was observed to 70 percent chord onboth surfaces at design conditions.

joints, rivets, and screw heads. Plaster splashes of the

upper and lower wing surfaces were made to measurewaviness. The measured waviness was welt below themaximum allowable for a single wave. (See Kosin1967.) Transitionwas realized to about 15 percentchord for 20 wing sweep and to about 5 percent chordfor 30 wing sweep. The amplification of TS disturbances is proposedto be the cause for transitionbecausetransitionwas realizedin the region of

A.7. F-14

adverse pressure gradient. The impact of engine noise

on transition was inconclusive. The flight test resultswere not compared

Following the achievement

of laminar flow on theF-Ill,the F-14 VariableSweep TransitionFlightExperiment(VSTFE)was initiated by NASA andBoeing CommercialAirplaneCompany (Anderson,Meyer, and Chiles 1988). Unlike the F-Illglove(which was not designed to minimize CF disturbancegrowth), the F-14 gloves were designed to optimizebetween TS- and CF-disturbancegrowth. The F-14

test used nearly all the span of the variable-sweep

portion and hot films to detect transition onset in the

determinethe potential influenceof sound on thepotentially unstable flows. Essentially,a less stableflow _'ould be expected simply by thickeningthenacelle lip. Obara and Dodbele (1987) reported theaerodynamicperformanceresults realized during theflight experimentand Schoensterand Jones (1987)reported the effect of the acoustic sources. For a flighttest at altitude of 1300 ft, Mach number of 0.25, and

A.8. NLF Nacelle

Flight

unit Reynolds number per foot of 1.8 x 106, subliming

chemicals indicated laminar flow to 50 percent of thenacellelength,with transitionoccurringat theforebody-aftbodyjoint. At the same flight conditions,the noise sources had no noticeable impact on the transition locations. Away from the pylon, the measuredpressure distributions were shown to be in good agreement with the design pressure back to the pressurepeak.

theA.9. Boeing

Experiment

757 NFL

The question

Flight

of whether

Testlaminar

flow could

be

maintainedon a commercialtransport with highbypass-ratiowing-mountedturbofan engines led toanother NASA-fundedflight experiment. The BoeingCompany used its Boeing 757 flight research aircraftwith a part of one wing modified to reduce sweep andobtain more NLF and to obtain extensive noise field

About the same time, a NLF nacelle flight experiment was conducted through a teaming effort led byGeneral Electric AircraftEngines. The experimentwas pursued because the friction drag associated withmodem turbofan nacelles may be as large as 4 to5 percent of the total aircraft drag for a typical commercial transport and because potential specific fuelconsumption(SFC) reductionson the order of 1 to1.5 percent may be achieved for laminar boundarylayer flows on advanced nacelles. The first phase ofthe flight experiment involved flying a NLF fairing onthe nacelle of a Citation aircraft to develop test techniques and to establish the feasibility of the concept.Hastings et al. (1986) reported the results of the firstphase which achieved laminar flow to 37 percent ofthe fairing length. The analysis showedthat theGranville (1953) criterion predicted the observed transition location for two of the four locations and that

measurementson a commercialtransport (Runyanet al. 1987). Primary goals of the experiment includedthe determinationof the influence of noise on the laminarwasThefaceand

boandary-layerflow. A 21 swept-wingglovemoanted outboard of the engine on the right wing.noise level was measured with microphones, surpressures were measured with strip-a-tube belts,transition locations with hot films as a function of

engine aower and flight condition. A large database

was obttined during the course of the flight test experiment. ?'he results suggest that the noise levels on thelower sarface have engine power dependence;however, the upper surface did not show engine powerdependencebut did show Mach number dependence.At the design point, laminar flow was observed to28 perc_:nt chord on the upper surface of the glove andto 18 pt:rcent chord on the lower surface. At the outboard portion of the glove, transition occurred at about5 percentchord where the pressure peaked(notpredicted by the transonic design code). The lowersurface was more sensitive to engine power and 2 to3 percent less laminar flow was observed at the higherpower settings compared with lower power settings.

the pressure on the fairing induced a neutrally stable

flow; this indicated that the flow was sensitive toexternal effects. The second phase of the flight testexperimentinvolved flying a full-scale flow-throughNLF nacelle (of various geometries) under the wing ofa Grumman OV-1 Mohawk aircraft (Hastings1987;Faust and Mungur 1987). Three nacelle shapes wereselected and designed to have pressure distributionswhich led to flow fields which were susceptibletoboundary-layerinstabilities.The variationwas to64

to 24 (obtained with sideslip). For a Mach number of

with the Boeing 757 and F-111

chord dependent on flap and yaw settings (Horstmann

et al. 1990). For TS-disturbance-dominatedtransition,the transition front was at nearly the same chordwise

flight database.A.10.

VFW

614

In 1986, the German

laminar flow technology

location across the span, whereas for CF-disturbancedominated transition, a distinct sawtooth pattern arose

pro-

(reminiscent of CF transition). As yaw was increased,

the laminar attachment line became intermittentlyturbulentwhich was consistentwith the threshold

gram, supported by the German Ministry of Research

and Technology(BMP-'I'), began wind tunnel andflight experimentsNLF and LFC (Redekeret al.1990). Korner (1990) noted that part of the programinvolveddetermining(or discriminating)betweenwhen NLF is preferred and when HLFC or LFC is amore appropriate choice for a particular aircraft. Additionally, two of the major milestones of this programinvolved NLF wind tunnel tests and flight research onthe 40-seat VFW 614 research aircraft (owned by

momentumReynolds number of 100 on the attachment line. Following the VFW 614 NLF flight test, aFokker 100 transport aircraft was fitted with a partialspan NLF glove to measure the drag reduction associated with a NLF wing design, validate laminar flow

DLR) during 1987 through 1990. The goal of the

VFW 614 ATTAS NLF flight experiment was to gaina database of TS-disturbanceand CF-disturbancedominated transition for code calibration. During the

CFD methodology,and to establish the upper limits ofNLF (transition Reynolds number for a given leadingedge sweep angle). The flight test consisted of threeflights for a total of 12 hr. The observedresultsvalidated the design predictions of 15-percent dragreduction;this confirmedhigh-speedwind tunnel

flight test, a database was obtained for variations in

In this section, a brief

NLF research is given.B.1. F-104

Starfighter

summary

Flight

wing. ".'he glove was 4 ft wide, extended past 30 percent chord, and a notch-bump(fig. 15) was added tothe inboard side of the leading edge of the test sectionto eliminate the potential for attachment-linecontamination problems. The flight tests were flown at Machnumbers ranging from 0.7 to 1.8, altitudes of 20000 to55 000 ft, unit Reynolds numbers per foot of 1.2 x 106to 4 x 106, and angles of attack of -1 to 10 . Com-

Flow

of supersonic

pressible stability calculations (using COSAL) for stationary crossflow disturbances at zero frequency werecorrelated with the flight-observedtransition location.Ignoring surface curvature, N-factors of 10.5 and 11matched the transition point for the Mach numbers of

Test

Some of the first transition-related

supersonicflight tests were carried out at the NASA High-SpeedFlightStation in California.In 1959, McTigue,Overton,and Petty (1959) reportedon transitiondetectiontechniquestested in supersonicflight byusing an F-104 Starfighter.A wing glove made offiberglass cloth and epoxy resin was positioned on thewing of the fighter-type aircraft. Resistance thermometers and subliming chemicals were used to detect thetransitionlocation.Cameraswere used to record

0.98 and 1.16, where the transition points were measured at 20 and 15 percent chord, respectively.Fortransitionoccurringcloserto the leadingedge,N-factors of 5.5 and 6 were found for Mach numbersof 0.9 and 1.76. Surfaceclean-upglovesweremounted on both the right wing (leading-edgesweepof 60 ) and the vertical tail (sweep of 55 ) of theF-106. Gaster-typebumpswere installedon theinboard portion of the gloves to prevent attachmentline contamination.Flight tests were conductedat

the sublimation process in flight. Approximately

40instrumentedflights were flown up to a Mach numberof 2.0 and an altitude of 55 000 ft. Photographs werepresented in the report giving a measure of transitionlocation (laminar flow extent) with various flight conditions. No detailed analysis of the transition locationand mean-flow attributes was performed.B.2. F-106

Mach numbers ranging from 0.8 to 1.8, altitudes ranging from 30 000 to 50 000 ft, unit Reynolds numbersper foot of 1.6 x 106 to 5.2 106, and angles of attackof 3 to 14 . Turbulent flow was observed at the firsthot-filrr gauge (0.5 percent chord) for all but four ofthe flig_at test points. All the transition points wereobserved within 5 percent chord of the leading edge.Either the attachment-linecontaminationpreventionwas not working properly or strong crossflow disturbances were generatedby the large leading-edgesweep. Collier and Johnson (1987) showed theoreti-

and F-15

An F-106 at Langley Research Center and an F-15

at Dryden Flight Research Center had a 6-month window of availabilityin 1985 which could be used tostudy supersonicboundary-layertransition(Collier

cally that N-factor

values could be significantlydecreasc'd by adding small quantities of suction in thefirst 12 percent chord of the vertical tail for a simulated F- 106 test point. With this small amount of suction, disturbances were stable to 20 percent chord; thissuggests, that HLFC would lead to significantrunswith lantinar flow.

and Johnson 1987). The F-15 twin-engine fighter was

selected as a flight test vehicle because earlier flighttests have shown that pressures on the 45 swept wingwould support small amounts of NLF. A surfaceclean-up glove was installed on the right wing of theF-15 to eliminate surface imperfectionsin the original

Groth, E. E. 1964a: Investigation

of a Laminar Flat PlateWith Suction Through Many Fine Slots With and Without Weak Incident Shock Waves. Summary of LaminarBoundaryLayerControlResearch.VolumeI,ASD-TDR-63-554,U.S. Air Force,pp. 428--441.(Available from DTIC as AD 605 185.)

Pearce, W. E. 1982: Progress at Douglas on Laminar Flow

ControlAppliedto CommercialTransportAircraft.ICAS Proceedings,1982: 13th Congress of the International Council of the AeronauticalSciences, AIAA Aircraft Systems and Technology Conference,B. Laschkaand R. Staufenbiel,eds.

suction. M > 1 ; h = 16.7 km;

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The history of Laminar Flow Control (LFC) from the 1930s through the 1990s is reviewed and the current status ofthe technology is assessed. Early studies related to the natural laminar boundary-layerflow physics, manufacturingtolerances for laminar flow, and insect-contaminationavoidance are discussed. Although most of this publication isabout slot-, porous-, and perforated-suctionLFC concept studies in wind tunnel and flight experiments,some mention is made of thermal LFC. Theoretical and computationaltools to describe the LFC aerodynamicsare includedfor completeness.